Scramjet Engine Research Articles

Effect of jet splitting using passive strut on the performance and thermoacoustic characteristics of a scramjet combustor

A scramjet engine offers a potential route to achieve supersonic speeds using airbreathing engines. Achieving proper mixing and combustion poses a challenge due to the supersonic inflow of air. Researchers have explored multi-strut configurations to tackle this issue. However, multiple struts supplying fuel inefficiently can lead to fuel loss and reduced efficiency. Alternatively, utilizing a multi-strut setup passively could enhance combustion and mixing efficiency. In this study, two types of jet splitting passive strut configurations were investigated computationally with the improved delayed detached-eddy simulation turbulence model. Implementation of passive strut altered vortical structures, influencing mixing and combustion performance. The splitting of the jet introduces large-scale vortices downstream. Strategically placing the passive strut in the wake of the combustion zone was found to improve both mixing and combustion efficiency. Acoustic loading was seen to increase with the introduction of passive strut. It was observed that the diamond-shaped passive strut has the highest combustion efficiency; however, it suffers from higher acoustic loading. The dynamic mode decomposition analysis revealed the coupling frequency of fluctuating pressure and heat release rate, which causes thermoacoustic loading. Overall, passive strut placement significantly influenced combustion, mixing, and thermoacoustic properties, highlighting the importance of considering passive strut configurations in design optimization for scramjet engines.

Effect of self-excited oscillation of three-dimensional rectangular shock-induced thrust vector nozzle jet

The fluidic thrust vector nozzles including the shock-induced thrust vector nozzles stand out from traditional mechanical thrust vector nozzles used in aeronautics and astronautics due to their simplicity and potential for higher efficiency. However, a significant challenge in the transition from theoretical studies to practical applications is the phenomenon of self-excited oscillation of the nozzle jet, particularly in ramjet and scramjet engines. This oscillation can notably impact the jet control stability, which is critical for the operational reliability, accuracy, and safety of these engines. To investigate the effects of self-excited oscillation of the jet in the three-dimensional rectangular shock-induced thrust vector nozzles, a large eddy simulation approach has been utilized to examine various nozzle pressure ratios and secondary pressure ratios. The simulation data are in good agreement with the experimental data of National Aeronautics and Space Administration Langley Research Center, lending credibility to the simulation results. The research sheds light on the formation and evolution of self-excited oscillation. It does so by examining the interactions between shock waves and boundary layers, as well as bubble dynamics, offering a comprehensive view of the oscillation mechanism in a three-dimensional context. The results demonstrate that the self-excited oscillation of jet mainly belongs to low-frequency oscillation. With the increase in nozzle pressure ratio, the self-excited oscillation of the jet is suppressed because the shock system is pushed out of the shock-induced thrust vector nozzle exit. The variation of secondary pressure ratio only affects the amplitude of jet self-excited oscillation and does not transform the motion pattern.

Study on the spray characteristics of transverse jets with elliptical nozzles in supersonic crossflows using the volume of fluid–discrete phase model

The atomization characteristics of liquid jets injected transversely into a supersonic crossflow significantly affect the performance of scramjet engines. Existing research on this topic has mainly focused on circular nozzles, while the influence of nozzle geometry, particularly elliptical nozzles, has received relatively limited attention. Therefore, this study employs a numerical simulation method coupling the volume of fluid and discrete particle model to investigate the breakup and atomization characteristics of transverse liquid jets from elliptical nozzles with different aspect ratios under supersonic crossflow conditions, as well as the total pressure loss. The simulation model is validated against experimental data obtained from a pulse wind tunnel, including measurements of the liquid jet penetration depth and the Sauter mean diameter (SMD). The results indicate that for elliptical nozzles with an aspect ratio (AR) greater than 1, columnar breakup occurs earlier, and the columnar breakup length is shorter compared to circular nozzles. As the AR increases, the jet penetration depth decreases, while the spray expansion angle increases. Furthermore, the SMD of the atomized spray field from the circular nozzle is larger than that from the elliptical nozzles, and the SMD of the spray field is smallest for an elliptical nozzle with AR of 4. Finally, the elliptical nozzles exhibit a higher total pressure recovery coefficient, indicating reduced total pressure loss in the combustion chamber. This reduction in pressure loss is expected to improve the thrust performance of the scramjet engine.

Performance analysis of a planar shaped strut injector based supersonic combustion chamber

Abstract This current work presents the performance analysis of a supersonic combustor interface and flow construction through a scramjet engine with a planar shaped strut injector (PSSI) at the supersonic Mach. An important aspect of this study is discovering the fuel mixing mechanism inside the combustion chamber with PSSI. The novel PSSI configuration can enhance mixing and combustion performance. The scramjet configuration is incorporated with a supersonic inlet air temperature of 1250 K, where the vitiated air follows at Mach 3, and this technique is based on a high accelerating effect of the scramjet propulsion mechanism. Scramjet engines can maintain naturalistic high enthalpy conditions in minimum durations. It is observed that the maximum temperature of 3510 K is attained at the recirculating zones produced because of undulation enlargement and therefore the fuel jet losses concentration whereas the maximum combustion efficiency of 86 % is investigated from the current research work.

An experimental investigation on transpiration cooling with phase change subjected to non-uniform flame

Transpiration cooling with phase change is promising for thermal protection of scramjet engine due to the combined advantages of latent heat absorption of coolant and thermal insulation of gas film. Previous investigations primarily focused on the transpiration cooling subjected to uniform incoming flow. However, few investigations have revealed the effect of the non-uniformity of high-temperature flame on transpiration cooling with phase change. Moreover, the effect of porosity on transpiration cooling performance was seldomly mentioned. Herein, we experimentally investigate the transpiration cooling in porous media of alloy (Inconel 600) with deionized water as the coolant subjected to the non-uniform high-temperature flame. Moreover, the effects of porosity and type of transpiration material on cooling are investigated. Results show that when the injection rate is less than 16 %, local overheating occurs on the surface of the porous materials. The overheating phenomenon occurs earlier at lower injection rates. Moreover, the decrease in porosity of the transpiration material will lead to increases of the surface temperature and top/back surface temperature difference, as well as decreases of top-surface temperature non-uniformity and cooling efficiency. In addition, for the different porous materials with similar thermal conductivity, the specific surface area exhibits a dominant factor in transpiration cooling.

Heat transfer enhancement mechanics of different micro-disturbing structures for transcritical hydrocarbon fuel using 3-D secondary flow intensity analysis

Heat transfer enhancement based upon micro-disturbing structures becomes increasingly important for thermal protection of air-breathing propulsion systems. To explore the influence mechanism of different structures on heat transfer in regenerative cooling channels of scramjet engines, flow structures and heat transfer of supercritical-pressure n-Decane in mini-channels with dimple and micro-rib arrays are numerically investigated. Results indicate effective heat transfer enhancement inside the channels depends more on efficient transverse turbulence disturbance rather than longitudinal or perpendicular turbulence disturbance. Severe thermal stratification generates strong shear layers in channels, where transverse secondary flows bring much intenser flow mixing than longitudinal secondary flows, especially in transcritical temperature region where the best heat transfer enhancement is obtained. With same characteristic dimensions, longitudinal micro-disturbing structures like dimples have lower flow resistance, but transverse micro-disturbing structures like micro-ribs achieve stronger heat transfer enhancement. Besides, the performance of micro-ribs is much more sensitive to the dimensional change. Therefore, micro-ribs have larger working excursion for enhancing performance factor than dimples. With optimal structure parameters: hr = e = 0.3 mm, pr = 3 mm, hd = 0.225 mm, dd = 1.5 mm, overall performance factor in micro-ribbed channel is 1.14 times of that in dimpled channel. Those conclusions provide effective support for optimization design for the regenerative cooling structures.

Refined mode classification and performance analysis method of dual-mode scramjet vehicles for flight trajectory optimization

Hypersonic air-breathing vehicles, utilizing ramjet and scramjet propulsion systems, are emerging as a promising option for future hypersonic transportation due to high specific impulse. The performance of hypersonic aircraft is significantly influenced by their flight trajectory, as combustion efficiency and aerodynamic forces vary markedly with altitude. Although numerous studies have focused on optimizing these trajectories, they have not adequately considered the potential for operational failures in the propulsion system, such as unstart and blowout. This study introduces a trajectory optimization approach for hypersonic aircraft that proactively addresses and mitigates these operational failures. This is achieved by establishing an operational classification process for a dual-mode scramjet engine and proposing a failure mode avoidance strategy using a Gaussian process classifier. Optimized ascent trajectories are achieved through the development of a two-stage robust sequential convex programming approach. The results of trajectory optimization indicate that the failure mode avoidance strategy proposed is crucial in obtaining minimal time trajectories, and that the optimal trajectories include drop motion in the initial stage flight in order to increase combustion efficiency.

Study of combustion flow field in a scramjet engine under inflow oscillation variations based on OpenFOAM

In this study, a compressible supersonic numerical model was established based on the open-source computational fluid dynamics software OpenFOAM. The supersonic air inflow was set in sinusoidal variation and gas hydrogen was employed as fuel for the DLR engine numerical simulation. The SST k-ω model and a 9-species 27-reaction hydrogen/oxygen chemical reaction model was employed as the turbulence and reaction mechanical kinetics models respectively. The effect of continuous variable inflow on the combustion performance for the scramjet engine was analyzed and the results showed that the combustion flame inside the scramjet engine exhibited significant partitioning phenomena around the combustion stabilization zones under periodic continuous variable inflow conditions. It was found that intense combustion zones were formed at both upstream and downstream flow as the induction of intersecting shock waves. When the flame was stabilized in the engine, the initiation position of the intense combustion zone at the upstream was transferred to the distance of twice the strut length, while the initiation position of the intense combustion zone at the downstream was anchored at a distance of about five times the strut length.

Convection Heat-Transfer Characteristics of Supercritical Pressure RP-3 in Horizontal Microchannels

To enhance the heat-transfer performance of scramjet engines, a numerical simulation was conducted on the heat-transfer process of RP-3 aviation kerosene under supercritical pressure within a horizontal micro-fine circular tube. The intrinsic mechanism of the heat-transfer process was analyzed, summarizing the impacts of mass flux, inlet temperature, and gravitational acceleration. Furthermore, four commonly used buoyancy criterion numbers were compared and evaluated. The results indicate that the heat-transfer process can be divided into five phases: heating inlet phase, normal heat-transfer phase, heat-transfer deterioration phase, heat-transfer enhancement phase, and high-temperature normal heat-transfer phase. The heating inlet phase is significantly influenced by the inlet temperature, while the heat-transfer deterioration is affected both by the thermal property variations of the aviation kerosene and the buoyancy effects. Lower mass flux and hypergravity conditions all exacerbate heat-transfer deterioration. Inlet temperature, however, does not affect the heat-transfer pattern. Among the criteria, Grq/Grth provides the best prediction of buoyancy effects in horizontal circular tubes.

Experimental investigation on the impact of fuel injection schemes on the fuel/air mixing characteristics in alternating wedge struts

To enhance the performance of scramjet engines, it is essential to understand the fuel/air mixing characteristics downstream of alternating wedge struts and the dynamics of streamwise vortices. Two optical measurement techniques, Mie scattering and acetone-particle laser-induced fluorescence, were employed to study the impact of fuel injection location and different injection gases on the fuel/air mixing characteristics downstream of alternating wedge struts at total temperature of 300 K and total pressure of 800 kPa. The results indicated that the particle laser-induced fluorescence measurements provided superior results compared to Mie scattering, due to the more uniform distribution and better following characteristics of acetone molecules as seeding particles. Different fuel injection locations resulted in varying fuel distributions and plume morphologies downstream. The plume sizes obtained from the three types of fuel orifices were generally consistent, but the mixing effect of the orifices on the first strut was better, whose injection locations being at the trailing edge intersections, with a larger plume area and more uniform fuel distribution. The fuel/air mixing process downstream of the alternating wedge struts was primarily dominated by streamwise vortices. Despite certain differences in the physical properties of nitrogen and hydrogen molecules, the measurement results showed that the plume morphologies obtained with both gases were quite similar. Additionally, a streamwise vortex model was developed to predict the dynamics of streamwise vortices and the evolution of fuel plume morphology, and it was compared with experimental results.

Ablation behavior at full-temperature ranges of C/SiC in combustion gases containing water vapor

In order to clarify the ablation behavior of C/SiC ceramics in the combustion gases containing water vapor, we proposed a novel ablation model of C/SiC at full-temperature ranges and simulated the ablation behavior of C/SiC ceramics under combustion gases environments by using our program developed in MATLAB. The numerical results indicate that the comparison between the calculation and the experimental data is in good agreement. The surface temperature of C/SiC ceramics is positively correlated with heat flux, enthalpy and pressure, while negatively correlated with the mass fraction of H2O in the combustion gases. Furthermore, the surface recession of C/SiC increases with the increase of heat flux or water vapor content but decreases with the increase of pressure. This study can predict the ablation behavior of C/SiC ceramics in combustion gases environments, and then guide for optimization of C/SiC components in scramjet engines.

Experimental Investigation of Microwave-Assisted Kerosene Supersonic Combustion

ABSTRACT The microwave-assisted kerosene combustion in supersonic flow was invested experimentally using a novel microwave-flame coupling method. The experiment used a direct-connected system, with a horn antenna employed to introduce microwaves into the combustion chamber, generating a traveling wave within the cavity. The microwave field simulation results confirm this coupling method’s feasibility. The experimental results showed that with the influence of the microwave, the flame’s stable position moved upstream by 7.4 mm. The average velocity of flame propagation increased by 64% upstream of entering the cavity and by 61% upon exiting it. The specific impulse of the scramjet was increased by about 4% under a 700W continuous microwave. In the presence of hydrogen, there was a significant increase in the excitation of NO and N2. However, in hydrogen-free combustion processes, the enhancement of radicals was considerably less compared to ignition processes, suggesting a close correlation between microwave-assisted combustion and hydrogen. These experimental results validate the feasibility of the microwave-combustion coupling method, representing a novel exploration into microwave-assisted kerosene combustion that aligns more closely with the practical application scenarios of scramjet engines.

Thermodynamic performance and flow structures analysis of supercritical CO2 applied in a regenerative cooling channel mounted with mini-ribs

Supercritical CO2 (SCO2) is regarded as a promising supplementary coolant for traditional regenerative cooling at extremely high flight speed (Ma ≥8) due to its superior heat and mass transfer capability. Using sCO2 as the coolant, mini-ribs is introduced to the rectangular regenerative cooling channel to solve the difficulties in transferring enough heat from near walls to core mainstream at a heat flux of 5 MW/m2 in this paper. With the thermophysical properties near the supercritical point interpolated in the solver, flow structures and heat transfer performance in the cooling channels are obtained using a validated k-ω SST turbulence model. Effects of pitch ratios, channel materials and acceleration and buoyancy effects are also investigated in terms of the local bulk heat transfer coefficients, Fanning friction factors and entropy generation numbers by heat transfer and fluid flow. Overall, the existing of mini-ribs improves the heat transfer coefficient and reduces the entropy number by heat transfer at the cost of the increased fluid friction and the corresponding entropy number. Compared with the smooth channel, the overall enhanced factor can reach 1.73 with an overall thermal performance factor of 1.43 and the maximum wall temperature decreases by 27.57 % in the case of P/e = 10. When the solid channel material has higher thermal conductivity, it leads to a more uniform temperature distribution and lower entropy generation number by heat transfer with the enhanced transverse heat conduction within solid walls which also leads large heat transfer coefficients. Larger accelerations result in stronger heat transfer and larger fluid friction by strengthening the buoyancy effect which is more obvious when the state approach the critical point. This work provides a good reference for the application of supercritical CO2 in a regenerative cooling channel of scramjet engine at extremely high heat flux.

Effect of perforated wall in controlling the separation due to SWBLI at Mach no. 5 to 9

Abstract The mixed compression scramjet engine intake was numerically simulated to study the characteristic of shock wave induced boundary layer separation leading to the formation of the separation bubble (SB). The analysis employed a 2D-RANS method with SST k-ω turbulence model at different Mach numbers. The intake is designed as a three-ramp intake for improved performance. As the Mach number varies, the size of the separation bubble formed due to the interaction of oblique shock waves with the boundary layer also varies, affecting both intake the efficiency and overall efficiency of the engine. Apart from traditional control techniques, the most preferred bleed technique is incorporated. Localised & distributed bleed techniques are designed and analysed within the intake at different locations. These techniques result in a reduction in the size of the separation bubble within the intake. Establishing perforation in the engine intake also increases the intake efficiency and overall engine performance.

Research on the Combustion Mode and Thrust Performance of Rotating Detonation Scrarmjet Engines

Decades of research in hypersonic science and engineering have shown that scramjet engines are the preferred choice for powering hypersonic vehicles. In these engines, oblique shock combustion and rotating detonation combustion are two common combustion modes. This article primarily focuses on the numerical simulation and theoretical analysis of these two combustion modes. The research indicates that as the combustible mixture injected into the combustion chamber increases, the combustion mode transitions from deflagration combustion to self-sustaining detonation combustion and then to forced detonation combustion to maintain stable combustion in the chamber. Under self-sustaining detonation combustion mode, as the amount of combustible mixture injected into the chamber increases, the number of detonation waves gradually increases from one to multiple and the angle between the detonation waves and the inflow changes from acute to right angles. In the forced detonation combustion mode, the minimum shock wave intensity required to initiate detonation combustion can be obtained by drawing a tangent from the Rayleigh line. The structure generating oblique shock waves in the combustion chamber needs to match the shock wave intensity. In the detonation combustion mode, both the exit total temperature and specific impulse of the combustion chamber increase with the increase in Mach number at the inlet, while the pressure ratio decreases. This article provides a reference for the design of the combustion chamber in supersonic combustion ramjet engines.

Influence of the non-equal aligned nozzles for fuel injection inside the supersonic combustion chamber

The importance of fuel mixing for the progress of the scramjet engine is indisputable. The present article shows the importance of the non-equal multi-injector system for effective fuel distribution and flame holding inside the combustion segment of a scramjet engine. The supersonic air and fuel jet flow in the non-equal nozzle arrangement is simulated via computational fluid dynamic technique. Two injector types of circular and rectangular nozzle have been analyzed to attain flow characteristics of hydrogen jets at supersonic cross flow. Mach contour is also analyzed for these jet arrangements to show the interface of the jet in the non-equal jet arrangement. Besides, addition of internal air jet is also simulated and evaluated in this research. Our investigation shows that the diffusion height of the fuel jet is higher when a rectangular non-equal nozzle is applied. The circular nozzle is more active in the spreading of the fuel in the combustor and the use of an internal air jet effectively increases fuel in a combustor of the scramjet.

Flow and heat transfer characteristics of hydrocarbon fuel in lightweight C/SiC regenerative cooling combustor

To meet the requirements of higher thermal protection ability and lower structural weight of regenerative cooling structures in advanced hypersonic vehicles, carbon-fiber-reinforced silicon carbide (C/SiC) composites with high permittable service temperatures and low densities are promising for application in thermal protection systems. In this study, four cylindrical regenerative cooling configurations applying C/SiC composites for a designated scramjet engine combustor are proposed, and the thermal protection performance under the influence of the cooling equivalence ratio (0.4∼1.0) and anisotropic thermal conductivity is studied. The results show that the C/SiC composites can significantly improve the heat resistance of the entire regenerative cooling structure owing to their relatively low thermal conductivity, which reduces the total heat flux transferred to the inner structure. The thermal protection ability can be maintained, whereas the coolant flow rate is reduced by 50%. C/SiC composites with higher thermal conductivities in the circumferential direction are beneficial for uniform temperature distribution. A flow distribution deterioration phenomenon caused by the thermal cracking of the coolant and the increasing flow resistance in the cooling channels can be observed under certain conditions, leading to a mass flow rate difference of 2.5 times among the cooling channels. Moreover, introducing C/SiC composites into the regenerative cooling structure design can significantly reduce the total structural weight by over 50% compared to traditional superalloys.

Enhancing fuel distribution of downward steps in supersonic combustion chambers using shock generators

This study investigates the impact of a shock generator on the fuel distribution of downward step fuel injection in a scramjet engine’s combustor. Computational simulations of CFD are employed to analyze the three-dimensional structure of multi-jets released in a downward-step configuration. The inlet is subjected to a free stream supersonic flow with [Formula: see text], while hydrogen gas is injected at [Formula: see text] through injectors situated on the vertical planes of the downward step. A shock generator with a [Formula: see text] angle is positioned at the top of the domain to induce a bow shock, which significantly affects fuel distribution. The research assesses the influence of free stream and jet pressure on the fuel distribution behind the jets. The results indicate that increasing the free stream flow strengthens the induced bow shock, leading to enhanced fuel jet interactions and improved fuel mixing behind the jets.

Supersonic combustion field evolution prediction in scramjet engine using a deblurring multi-scale attention network

Traditional deep learning technologies often overlook crucial spatiotemporal information when reconstructing supersonic combustion flow fields. These models can only reconstruct the current flow field, resulting in poor visual quality due to the atomization effect. This research introduces experimental studies on scramjet combustion flows with variable injection pressures at the inlet of a scramjet Mach 2.5, creating image datasets of flame evolution processes over different predicted time spans (5, 10, and 15 ms). It develops a multi-scale attention algorithm with deblurring (DB-MSAN) capabilities to predict the flame image after a certain period using the wall pressure signal from the previous moment. The algorithm employs two multi-head attention channels at different scales to extract timing information, encouraging the model to focus on the anisotropy of the input data and to identify more significant features. The DB module utilizes the adapted atmospheric scattering model to jointly estimate the transmission map and atmospheric light, achieving an exceptional defogging effect with a minimal computational expense. The model’s overall performance is assessed regarding efficiency and accuracy, comparing DB-MSAN’s performance differences with those of Chen’s and the ResNet16 models across multi-span, large-span, various test sets, and sparse wall measurement points. Experimental results indicated that, compared to other methods, DB-MSAN achieves a maximum peak signal-to-noise index improvement of 76.10 % and an SSIM index improvement of 155.26 % when predicting average-quality images. The prediction accuracy remains stable, even with sparse pressure input. In addition, DB-MSAN strikes an optimal balance between the number of parameters and model accuracy; it represents only 16.97 % of the volume compared to Chen’s model, which is 585 MB.

Parametric Study of Flow and Combustion Characteristic in a Cavitied Scramjet with Multi-Position Injection

This study focuses on the three-dimensional flow and combustion characteristics of a cavitied scramjet engine with multi-position injection. A single-equation large eddy simulation (LES) turbulence model is employed, with a detailed reaction mechanism for hydrogen combustion, as described by Jachimowski. The combustion characteristics of hydrogen in the scramjet combustion chamber are analyzed. Based on the combustion chamber model, the influence of different equivalence ratios, injection timing, injection positions, and injection pressures on the flame formation and propagation process are compared. The results indicate that within a certain range, an increase in the equivalence ratio enhances the combustion intensity and chamber pressure. In the case of multi-position injection, the order of injection from different nozzles has little effect on the final flame stabilization mode and pressure distribution. The opposite-side distribution of nozzles can effectively improve the fuel efficiency and the internal pressure. Furthermore, when the nozzles are closely placed in the opposite-side distribution, the combustion efficiency increases, although this leads to a higher total pressure loss. In scenarios where the fuel injection duration is short, an increase in the injection pressure at the upstream nozzles of the cavity results in a higher local equivalence ratio, as well as reduced fuel mixing and ignition time.