Scramjet Engine Research Articles

Study on combined thermal protection scheme integrating supercritical CO2 regenerative cooling and fuel film cooling used for scramjet engines

Supercritical CO2 closed-Brayton-cycle demonstrates considerable application potential on hypersonic vehicles, effectively addressing power supply issues during prolonged flights. However, the CO2 flow rate is strictly limited by the cycle’s cooling source (aviation kerosene), rendering traditional regenerative cooling inadequate for sufficiently cooling the scramjet engine. To address this challenge, this study proposes an integrated cooling scheme that combines regenerative cooling, ceramic thermal-protective layer, and fuel film cooling. A two-dimensional coupled numerical model was developed, simultaneously considering the convective heat transfer of supercritical CO2 and the shear mixing of supersonic fuel film. The results indicate that the high-enthalpy gas imposes an extreme aerodynamic heat load on the regenerative cooling channel wall, resulting in a notable deterioration in heat transfer for supercritical CO2. Implementing kerosene film cooling within the combustor can effectively mitigate the abnormal heat transfer behavior of supercritical CO2 and significantly reduce viscous dissipation heat within the gas boundary layer. By optimizing the fuel film injection layout, this combined cooling scheme reduced wall heat flux by 43% and temperature by over 400 K compared to traditional regenerative cooling, thus providing effective thermal protection for the scramjet engine.

Instability of isolator shocks to fuel flow rate modulations in a strut-stabilised scramjet combustor

Abstract Reynolds-Averaged Navier–Stokes (RANS) simulations, both steady and unsteady, are used to investigate supersonic, chemically reacting, flow fields inside a strut-stabilised supersonic combustion ramjet (scramjet) engine operating under different fuel flow rates. Fully supersonic, fully subsonic and mixed modes of operations inside the combustor, obtained at different fuel flow rates, are studied numerically through shock wave visualisations and top-wall static-pressure probing. The effect of changing fuel flow rates, imposed both suddenly and gradually, on the behaviour of shock waves and wall pressure profiles are studied in detail. For certain modes of combustion characterised by the presence of oblique shocks at the strut, shockwaves in the combustor respond predictably to an increase or decrease in fuel flow rate attaining the steady state flow fields as predicted by RANS simulations for those fuel flow rates. For certain other modes of combustion, characterised by the presence of shockwaves in the isolator and the absence of oblique shocks at the leading edge of the strut, shockwaves in the flow field appear unstable to fuel flow rate modulations. For such cases, any change in fuel flow rates, sudden or gradual, increase or decrease, causes the isolator shocks to immediately move upstream and eventually out of the isolator. A plausible physics-based explanation of the observed phenomena is presented.

Performance evaluation of a scramjet engine utilizing varied cavity aft wall divergence with parallel injection in a reacting flow field

Abstract The effect of implications of the dual cavity with aft wall divergence in a parallel injection reacting flow has been numerically investigated. This research intends to emphasize the behaviour of the supersonic flow under varying divergence angles of the cavity aft wall. A two-dimensional Reynolds-Averaged Navier-Stokes (RANS) equation and an SST k-ω turbulence model with a single-step chemical reaction for hydrogen-air are utilized for the simulation. Followed by the strut injector, the cavities are positioned symmetrically inside the combustor. The bottom cavity aft wall divergence varies, whereas the top wall is mounted with a rectangular cavity. The performance of cavity locations is compared with the baseline DLR model. From the evaluation of numerical outcomes along with different cavity configurations, it has been noted that the 15-degree cavity divergence angle enhances the recirculation zone which leads to improved mixing performance. Also, the cavity improves the combustion stability by increasing the flow residence time. From this numerical analysis, it is associated that an almost 20 % reduction in combustor length is achieved, however, an 18 % rise in the pressure loss is noted because of emanating shock waves from the cavity edges.

Study on solid powder fuel transport characteristics and mixing degree in scramjet combustor

Abstract The powder scramjet engine is considered the best engine for hypersonic weapons because of its advantages of easy adjustment, flame stability, and high efficiency. However, due to the extremely short residence time of particle fuel in the combustion chamber under high Mach conditions, there is a risk of fuel mixing. Based on the existing combustion chamber models of scramjet engines, this paper proposes a new solid fuel blending enhancement structure by combining shock generator structure and transverse jet and compares the performance and mass transfer characteristics of different shock generators. The CFD method is used to solve the RANS formula to simulate the particle transfer process, and the SST k-ω model is used to calculate the complex turbulence. It was found that when the reverse shock generator was used, the mixing degree of solid particles was higher.

Energy and exergy analyses on the high-pressure bleeding-air variable cycle designed for a wide-speed-range airbreathing propulsion system

The performance of the traditional turbojet decreases sharply with the increase of the Mach number in the supersonic flight, especially for the turbine-based combined cycle (TBCC) engine in the transition region of the turbojet and scramjet engine. This paper proposes a new high-pressure bleeding-air variable cycle engine (HB-VCE) concept that aims to ease the sharp contradiction between thrust output and fuel consumption of TBCC in the transition region. Through the energy and exergy analyses, the HB-VCE shows greater potential in energy utilization under high Mach flight conditions. The performance of HB-VCE significantly improves as the bleed ratio increases, with a more pronounced effect at higher Mach numbers. At Mach 1.5, the HB-VCE increases the specific thrust (ST) by 2.84 % and reduces the specific fuel consumption (SFC) by 2.73 %. At Mach 2.4, it increases the ST by 6.60 % and reduces the SFC by 6.52 %. Additionally, the controlling law of the critical bleeding ratio constrained by the turbine outlet temperature is proposed. Finally, the effect of the turbine inlet temperature, overall pressure ratio, and Mach number on the critical bleeding ratio is analyzed. This paper provides a promising HB-VCE concept with greater energy utilization for the TBCC design in the transition region.

Investigation of unswept ramp system with lobe shape nozzle for fuel mixing of hydrogen jet at a scramjet engine

The role of efficient fuel mixing and a stable flame holder is crucial in enhancing the performance and capabilities of scramjet engines for high-speed flight. The present research paper has tried to disclose the fuel mixing efficiency of 3-lobe annular nozzle on the mixing mechanism of the fuel jet behind the strut. In addition, using internal air jet flow for increasing the circulation strength and fuel mixing behind the strut is also examined in this study. Numerical simulation of the flow and fuel jet behind the strut is done to reveal the main physics related to the mechanism of fuel mixing inside the combustor with the proposed injection system. The results of our simulation show that using annular 3-lobe fuel jet improve the fuel mixing via production of the multiple vortex pairs within the combustor behind the strut. The use of internal air jet also enhances the fuel mixing efficiency up to 90% in combustor of scramjet engine.

Research on combustion stability characteristics of scramjet combustor equipped with strut injector

The combustor is the core component of the scramjet engine, and stable combustion is the prerequisite for the efficient operation of the scramjet engine. At present, researchers premix the fuel and oxidant to improve the combustion performance of the combustor, thereby reducing the mixing time and improving the combustion efficiency. This study proposes a new method of supplementing oxygen at the trailing edge of the strut to enhance mixing and combustion. This paper studies the combustion stability of the combustor equipped with this strut. This paper conducts a series of large eddy simulations under the combustor inlet conditions of Ma = 2.8, Tt = 1680 K, and Pt = 1.87 MPa, and obtains non-reactive flow field and reactive flow field data. The results show that the injection of oxygen at the trailing edge of the strut has a significant positive effect on fuel mixing and combustion. The non-reactive flow field results show that the injection of oxygen dilutes the kerosene at the trailing edge of the strut and improves the mixing efficiency. The reactive flow field results show that only when the oxygen supplement mass ratio is above 0.03 can a stable overall flame be formed inside the combustor, and oxygen supplementation enhances the heat release of premixed combustion. The Ret-Da distribution shows that due to the good mixing at the trailing edge of the strut, mixing is no longer the reason that restricts the combustion process, the heat release increases significantly, and the high-temperature heat release zone at the trailing edge of the strut ignites other fuels to form a stable flame in the combustor.

Research on the heat transfer characteristics of through-hole transpiration cooling based on ceramic matrix composites

Ceramic matrix composites (CMCs) are one of the primary materials used for transpiration cooling in aircraft due to their lightweight and high-temperature durability characteristics. However, there is relatively little research on the transpiration cooling of through-holes in CMCs at high temperatures. This paper establishes a numerical model for flow and heat transfer of through-holes transpiration cooling in the combustion chamber environment of a scramjet engine, analyzing the effects of blowing ratio, hole arrangement, hole size, and material thermal conductivity on the effectiveness of transpiration cooling, revealing potential strategies for optimizing hole arrangement using Response Surface Methodology (RSM). The calculation results show that increasing the blowing ratio can effectively enhance transpiration cooling performance. However, a higher blowing ratio increases the pressure gradient within the through-holes, leading to non-uniform coolant distribution and an increased temperature gradient. The plum hole arrangement provides better cooling performance than the parallel hole arrangement. Under the same blowing ratio, cooling efficiency improves by 8%, and uniformity increases by 50%. Increasing the size of the leading-edge holes can effectively reduce the pressure gradient within the through-holes. After optimization, the temperature gradient and average temperature improve by 8.824% and 4.290%, respectively. Additionally, enhancing the thermal conductivity in the direction of heat flow increases cooling efficiency more significantly compared to other directions.

Wavelet optical flow velocimetry of a scramjet combustor using high-speed frame-straddling focusing schlieren images

Scramjet combustors feature ultra-high-speed turbulent reacting flows with high temperatures and intense luminescence. Measurement of velocity fields under such extreme conditions presents great challenges. The present work demonstrated a seedless velocimetry approach using focusing schlieren images (FSIs) of high spatiotemporal resolution in a scramjet engine. Two fuel mass flow rates (Case1 and Case2) were investigated with corresponding global equivalence ratios of 0.27 and 0.13, respectively. The FSIs enabled by the employment of a high-speed pulsed LED light source are characterized by an effective exposure of 100 ns, and a 500-ns frame-straddling time interval with full resolution of 1280 × 800 pixels recorded at 76 kHz. The 100-ns exposure allows for capturing of transient high-speed flow motion without blurring, and the 500-ns time interval ensures an appropriate spatiotemporal correlation between subsequent schlieren images for high-speed reacting flows. A wavelet-based optical flow velocimetry (wOFV) algorithm was developed and applied to the FSIs. In contrast to the correlation-based algorithms widely employed in PIV for distinct particles, the wOFV algorithm suits better FSIs with continuous variation in brightness. The maximal velocities in the main duct of the scramjet combustor were measured to be approximately 550 m s-1 and 1100 m s-1 for Case 1 and Case2, suggestively corresponding to subsonic and supersonic combustion modes, respectively. The measured velocity inside the cavity is generally below 200 m s-1 for both cases. Recirculation regions and their dynamic motions inside the cavity were well resolved. In summary, the development of the novel velocimetry approach holds great potential for applications in extreme flow conditions. Novelty and Significance Statement: Present work demonstrates a seedless velocimetry approach based on high-speed frame-straddling focusing schlieren imaging coupled with the novel wavelet optical flow velocimetry (wOFV) algorithm. Velocity field measurement realized in a scramjet combustor with high spatiotemporal resolution (1280 × 800 pixels at 38 kHz) for the first time, showing great abilities to accommodate wide velocity range and to resolve dynamic flow characteristics with potentials for broader future applications.

Fuel mixing and diffusion behind the unswept ramp injector with lobed nozzle at combustor of scramjet engine

-The use of 2-lobe nozzle for efficient fuel mixing behind the ramp injector nozzle is comprehensively investigated in this article. Computational fluid dynamic is applied for the evaluation of the three-dimensional flow behind the strut with annular 2-lobe nozzle at supersonic combustion chamber. In this work, the Mach number of free steam air flow and fuel jet are 2 and 1, respectively. The usage of internal air jet for the fuel mixing improvement is also analyzed to disclose the critical features which improve the fuel diffusion in the combustor. The Mach contour on the horizontal and vertical planes in both injection model is compared to show how induced vortex pair is created and extended behind the strut. The result of computational study shows that the injection of the air jet increases the number of vortex pairs behind the strut and the interaction of the fuel with air jet is increased. However, the circulation strength is decreased by inner air jet. Use of inner jet enhances the fuel mixing more than two times behind the nozzle. Our findings show that the fuel mixing is heightened downstream of strut by coupling inner air jet with annular fuel jet.

Prediction model for self-starting of hypersonic inlets with soft critical unstart mode

The acceleration self-starting performance of hypersonic inlets is of critical importance for the stable operation of scramjet engines. The occurrence of soft unstart during the transition from hard unstart to start is an important flow state that has yet to be fully elucidated. The stability mechanism and corresponding self-starting characteristics of soft unstart remain poorly understood, and there is a pressing need for detailed modeling research in this area. This paper presents a rapid prediction model for the self-starting Mach number of two-dimensional hypersonic inlets with soft critical unstart mode, fully considering the influence of various geometric parameters and Reynolds number in the internal contraction section, and achieving a quantitative analysis of the two-dimensional soft unstart critical flow field. Given the incoming flow conditions and the inlet geometry, the prediction model is capable of accurately representing the actual viscous unstart flow field. It can fully map the unstart separation bubble and its surrounding critical wave structures, and calculate the minimum pressure rise required to maintain the current scale of the main separation bubble and the pressure rise exerted on the unstart separation bubble by the current actual flow field structure. Comparing the relative magnitude of these two pressures determines whether the inlet can transition from soft unstart to start. The proposed prediction model was validated using results from unsteady numerical simulations. The predicted results align well with the simulation results and are significantly better than previous prediction methods.

Chemical timescale analysis of the Partially Stirred Reactor model for a hydrogen-fuelled scramjet

Scramjet engines offer a promising option for powering hypersonic vehicles, although the unique characteristics of supersonic combustion present numerous challenges. Notably, turbulence-chemistry interactions are significant in supersonic flows, demanding specific combustion formulations in numerical simulations due to their dissimilarity from traditional combustion systems. This work evaluates the performance of the Partially Stirred Reactor (PaSR) closure of the mean chemical source term by performing Reynolds-Averaged Navier-Stokes (RANS) simulations of the German Aerospace Centre (DLR) scramjet. The PaSR approach requires defining a chemical timescale, and this study assesses several formulations. Among them, the Fastest Formation Rate (FFR) does not provide stable combustion, while the Slowest Formation Rate (SFR) achieve fairly good predictions. Moreover, the sensitivity to species selection in calculating the chemical timescale is investigated, as well as the impact of two PaSR residence time formulations. A newly proposed Batchelor-based mixing timescale performs well with both the SFR and FFR formulations and guarantees a coupling between chemical and mixing times. It also accounts for species diffusivity, which can have important effects on hydrogen-based fuel combustion.

Research on the flame oscillation characteristics of scramjet combustor equipped with strut fuel injector

The combustor is the core component of the scramjet engine, and stable combustion is the prerequisite for the efficient operation of the scramjet engine. This paper investigates combustion stability in a combustor with a thin strut for oxygen supplementation at the trailing edge. LES (Large eddy simulation) was carried out under the combustor inlet conditions of Ma= 2.8, Tt= 1680 K, and Pt= 1.87 MPa. The results show that as the amount of oxygen supplementation inside the combustor increases, the oscillation forward propagation characteristics of the flame front inside the combustor are formed. First, the combustion instability phenomenon of the flame inside the combustor is analyzed, and then the oscillation frequency of the flame is analyzed by FFT (Fast Fourier Transform) of the flame front and pressure and DMD (Dynamic Mode Decomposition) of the velocity field within one cycle. Finally, the flame forward propagation mechanism inside the combustor is analyzed by pressure-velocity-heat release distribution. The results show that the injection of excess oxygen inside the combustor forms a strong coupling phenomenon of flow and combustion, and the premixed gas downstream of the trailing edge of the strut forms a state of slow combustion to explosion, forming a flame propagation forward and periodic oscillation phenomenon.

Experimental investigation on cooling characteristic of transpiration-film combined cooling structure using liquid coolant

Transpiration cooling has gained significant attention as a promising strategy for thermal protection in scramjet engines. However, this approach faces limitations due to its susceptible block in fuel gas environment and significant fluctuation when using liquid coolants. To address these challenges, this work proposes a novel transpiration-film combined cooling structure (CCS), where the lower porous layer could ensure sufficient heat transfer and the upper film-hole layer could keep impurities from blocking porous layer and enhance the mechanical strength. Experimental and theoretical investigation on the combined cooling structure with liquid coolant phase change are carried out, and the cooling efficiency, temperature uniformity and cooling chamber pressure of the combined structure are compared with those of the transpiration cooling structure (TCS). Remarkably, the combined structure effectively suppresses temperature fluctuations, reducing the amplitude of temperature variation to just 1/5 of that observed in the transpiration cooling system. Although the combined structure shows a modest 0.026 drop in time-averaged cooling efficiency, it significantly improves temperature uniformity and reduces pressure peaks within the cooling chamber. Furthermore, the flow characteristics of liquid film and vapor film are analyzed and the coolant utilization rate for vapor film is 6.4 times higher than that for liquid film. Consequently, the design of combined structures using liquid coolant holds great promise for achieving highly efficient thermal protection in scramjet engines.

Optimization of Thrust of a Generic X-51 Hypersonic Vehicle

A prune-and-search optimization algorithm was combined with a reduced-order model of the scramjet engine within a generic X-51 vehicle. The goal was to maximize the thrust in order to achieve thrust-to-drag ratios that equal or exceed 1, even for the difficult cases of flight at high altitudes and high Mach numbers up to 10. For the inlet to the engine, the lengths and inclination angles of three wall panels were varied. For the combustor, three parameters were varied: the diameter and number of fuel injector ports and the wall divergence angle. These variables affect both the combustion efficiency and the finite-rate chemistry of the surrogate JP-7 fuel. The optimization algorithm showed that at higher flight Mach numbers an inlet that maximizes static pressure ratio is preferred over one that maximizes the stagnation pressure ratio, because good combustion efficiency requires a sufficient static pressure in the combustor. As the flight Mach number is increased, thrust decreases while the drag increases. Thus, it becomes difficult to achieve thrust that exceeds drag. A solution is shown to be increasing the engine inlet area above a critical value that was computed. Advantages of a reduced-order model over a high-fidelity CFD approach are discussed when thousands of computations of all components within an entire engine are required for optimization.

Pressure feedback system for flow separation mitigation in scramjet intakes

Shock Wave Boundary Layer Interactions (SWBLI) occur due to the convergence of a shock wave and a viscous boundary layer. The resulting separation bubble is a major setback for the performance of high-speed air intakes. This paper presents a numerical investigation into the potential of a Pressure Feedback Technique (PFT) to alleviate shock-induced flow separation within a Scramjet engine intake operating at Mach 4. The PFT is a novel self-sustaining flow control technique that combines simultaneous suction and injection. The two main output parameters used to support the hypothesis of the present study are the size of the separation bubble as well as the total pressure recovery at the isolator outlet. The most prominent observation of this study is that with the installation of the PF tubes, the total pressure recovery at the exit is noted to rise. The effectiveness of the pressure feedback technique in reducing the intensity of the separation bubbles is found to depend on the diameter of the PFT, pitch-to-diameter ratio, and PFT tube design.

Flow Effects and Propulsion Performance on Various Single Expansion Ramp Nozzle Configurations of Scramjet Engine

This study serves as a research endeavor aiming to explore the behavior of the coupling flow effects of the single expansion ramp nozzle (SERN) in over-expansion conditions during the static start-up process. The open-source program OpenFOAM and its solver “rhoCentralFoam” are employed in the 2D simulation and the two critical geometric variations, the shape of the ramp and the length of the flap beyond the throat, are considered in the geometric variation. The result shows the preferable propulsion performance in the FSS (Free Shockwave Separation) state compared to RSS (Restricted Shockwave Separation). FSS also plays the role of the initial, albeit transient, separation, which originates from the shockwave from the throat and will eventually transform into a stabler RSS state. For the 100% flap length configuration in this study, the axial thrust can achieve a high value of 500 N/m in the FSS state and decrease to around 450 N/m, on average, in the RSS state. The trust angle also shows a preferable performance of around −13° in FSS compared to −30° in RSS. Regarding geometric modifications, both modifications, shorting the flap and bell-shaped ramp adjustments, manifest similar effects. Both conical and bell-shaped short flap configurations demonstrate an axial thrust from around 1750 to 1900 N/m and a thrust angle of around −45°. However, the flap shortening, which may demonstrate an attitude compensation effect, exhibits a more pronounced effect compared to the bell-shaped modification.

The fuel mixing of multi-annular extruded fuel jets in cavity flame holder at supersonic combustion chamber: Computational study

A cavity-flame holder is widely used for the injection of the fuel jet in a supersonic combustor due to high fuel residence time. This paper presents extensive results about the impacts of extruded annular multi-jets for improving fuel mixing inside the cavity flame holder. A three-dimensional model of the cavity with multiple extruded nozzles in different lengths and conditions has been examined in this study to investigate the flow and fuel mixing inside the cavity. In addition, the usage of internal air jets is investigated and extensive comparisons between the elongation of the extruded nozzle and the use of the inner air jets for better fuel distribution are done. The presented results show that the usage of both inner air jet and nozzle elongation significantly enhances the fuel distribution within the cavity flame-holder. The achieved finding confirms that the latter technique is more efficient for the fuel mixing in the combustor of a scramjet engine.

Thermodynamic cycle performance modeling and numerical simulation of higher Mach scramjet with inlet pre-injection

This paper focuses on the thermodynamic cycle performance modeling and numerical investigation of a high Mach number scramjet with inlet pre-injection of hydrogen. The scramjet model used in this study was the Hyshot-Ⅱ engine model. Flight experiments using this model at high Mach number have been conducted previously, which provides a unique testbed for validating the computational prediction of supersonic combustion. A thermodynamic cycle performance model with inlet pre-injection was developed for operating characteristics studies of the scramjet engine. A set of three-dimensional Reynolds average Navier-Stockes (RANS) simulations of the reactive flow was performed with a 9-specie and 19-step kinetic mechanism for hydrogen combustion. The turbulence is modeled by k-ω shear stress transport (SST) turbulence model. The thermodynamic cycle performance analysis results indicate that inlet pre-injection can enhance the hydrogen mixing efficiency, and a small amount of advanced heat release can improve the overall thermal cycle efficiency. Detailed flow analysis from the three-dimensional simulation results indicates flow can be ignited in the first shock wave induced boundary layer separation (SIBLS) region. Inlet pre-injection enables the propagation of the flame toward the core flow compared to the throat injection. The inlet fuel pre-injection proved to be a promising injection method due to the higher combustion efficiency, elevated from 0.75 to 0.85, and higher specific impulse potential, increased from 1570 s to 1810 s. As the ER of inlet pre-injection increases within the examined range, the performance curve exhibits an optimal point. As the geometric compression ratio increases, the entire combustion process advances, leading to a significant increase in the performance potential of the engine. The comparative analysis of the thermodynamic cycle performance modeling and computational fluid dynamics (CFD) modeling indicates the model shows a good prediction on the performance of the scramjet engine with the inlet pre-injection, with the same trend of performance curves and an error of less than 10 % compared to the numerical simulation.

Comparison of fluid structure downstream of an evaporative flameholder with subsonic uniform and subsonic-supersonic inflows

Subsonic uniform and supersonic mixing flows are widely present in the combustion chamber of a multi-mode combined scramjet engine, where the significant pressure, velocity, and temperature gradients of the two impose severe limitations on flame propagation. The vortex structures downstream of an evaporative flameholder were experimentally measured under subsonic-supersonic mixing inflow and subsonic uniform inflow. Results indicate that, under subsonic-supersonic mixing inflow conditions, an asymmetric vortex structure in the recirculation zone is formed due to the presence of a subsonic-supersonic shear layer, which exerts compressive or stretching effects on the recirculation zone downstream of the flameholder, in contrast to the symmetric dual-vortex structure observed under subsonic uniform inflow conditions. The asymmetric recirculation zone displays fluid being entrained from one vortex to another, with the degree and direction of entrainment dependent on the supersonic expansion state. Additionally, a simplified flow field structure is proposed for comparing the differences between the subsonic-supersonic mixing inflow and subsonic uniform inflow, along with the introduction of five regions and two stages to describe the flow field structure downstream of the flameholder under subsonic-supersonic mixing inflow. Analysis of the flow characteristics downstream of the flameholder under subsonic-supersonic mixing inflow and subsonic uniform inflow conditions can offer design insights for the combustion scheme of aerospace-integrated propulsion systems.