Scramjet Engine Research Articles

Reacting flow analysis in scramjet engine: effect of mass flow rate of fuel and flight velocity

Abstract The objective of the study is realizing the effect of fuel mass flow rate and flight speed on combustion in scramjet engine. DLR conical strut based scramjet combustor configuration was chosen and simulated the chemical reaction between the air and hydrogen fuel. A slot of size 40 mm × 0.295 mm provided at the center of the strut to injected hydrogen fuel from the rare side in to the downstream flow. ICEM CFD software is used for the generation of structured elements in computational domain for three dimensional flow analyses. Standard k-epsilon turbulence model and species transport equation is used in ANSYS fluent solver. The predicted temperature, velocity distribution along the axial length was compared with the experimental results and validated. The temperature distribution at different Mach numbers and mass flow rate reveals that the peak temperature increased with the flight speed and inlet fuel mass flow rate. The peak temperature noticed at the center of the combustor is around 3500 K at a flight speed of Mach 4. The predicted variation of temperature, pressure, velocity in the combustor and the flow structure for reacting flow facilitate good understanding of the combustion process in scramjet combustor.

Combustion dynamics of high Mach number scramjet under different inflow thermal nonequilibrium conditions

To characterize the impact of thermal nonequilibrium conditions in shock-induced combustion in high Mach number scramjet engines, a comparative numerical investigation has been performed. The combustor configuration was based on the HyShot II scramjet while the total enthalpy is 3.3 MJ/kg which resembles Mach 8 flight condition. The improved delay detached eddy simulation with Park two-temperature approach in combination with the vibration-chemical coupling model was employed in the nonequilibrium simulations. It is shown that under thermal nonequilibrium inflow conditions, noticeable stratification in the thermal state exists in the circumferential direction both in the combustor section and exhaust nozzle. In the inner core where the fuel mixes with the mainstream, flow deceleration caused by the heat release leads to a sufficient transformation of energy between the vibrational and translational-rotational modes approaching the thermal equilibrium state. However, the τv,mix is quite large in the supersonic mainstream, whereby the thermal nonequilibrium dominates. For the case with higher Tv in the inflow, the fuel jet's penetration height is lower while the local viscosity is higher which results in poor mixing between the fuel and air stream. Nevertheless, a 400 K increase in inflow Tv gives a higher effective temperature, which promotes the dissociation reactions, shortens the ignition distance and yields an 8% increase in the overall combustion efficiency. With better combustion performance, an additional 6% deceleration in the supersonic mainstream and the energy exchange rate between vibrational and translational-rotational modes becomes almost four times faster at the combustor outlet, which promotes the restoration of the thermal equilibrium state. Furthermore, under higher inflow vibrational temperature, the shock-induced combustion in supersonic crossflow becomes less stable with noticeable fluctuations in combustion heat release and local static pressure.

Scramjet Engine Flowpath That Improves Specific Impulse Using JP-7 Fuel

A flowpath geometry was computed that improves the specific impulse of a dual-mode scramjet engine in a generic X-51 hypersonic vehicle. Six parameters were varied: inlet contraction ratio, the diameters and numbers of fuel injectors, divergence angle of the combustor wall, nozzle flap deflection angle, and flight Mach number. The maximum specific impulse was 2296 s in the ram mode and 832 s in the scram mode. A reduced-order model simulates the finite-rate chemistry of JP-7 fuel and the unstart limits. Results show that both combustion efficiency and drop to unacceptably low levels when the finite-rate chemical reaction rates are weakened by flame strain-out due to the large air velocities, or when the flame becomes longer than the combustor. Small occurs when the following are too small: the inlet contraction ratio, the inlet compression ratio, the number of fuel injectors, and the diameter of fuel injectors. When these parameters are too large, excessive heat release causes unstart. The operating range was identified between these limits. For JP-7 fuel, it was found that the inlet should raise the pressure to above 5 atm. Results are explained by the interactions between reactions, mixing, and flame strain-out.

Fast and reliable prediction of scramjet flowfields via Gaussian process latent variable model and deep learning

Fast and accurate prediction of high-speed flowfields is of particular interest to researchers in fluid science and engineering to enable efficient design exploration and knowledge discovery. The reliability of prediction is another important metric for the performance of prediction models. While predictive modeling approaches with and without reduced-order modeling (ROM) via machine learning techniques have been proposed, they are inherently subject to loss of information for ROM-based approaches and substantial computational costs in modeling for non-ROM-based approaches. This paper proposes an accurate ROM-based predictive framework with minimum information loss enabled by incorporating Gaussian process latent variable modeling (GPLVM) and deep learning. The stochastic nature of GPLVM allows for uncertainty quantification that indicates the degree of prediction error or reliability of prediction without requiring validation data. The applicability for supersonic/hypersonic viscous flowfields has been examined for two cases including axisymmetric intakes and two-dimensional fuel injection in scramjet engines by comparison with other predictive models. Comparable or superior prediction accuracy over the other models has been achieved by the proposed approaches, demonstrating its high potential to serve as a new competent, data-driven technique for fast, accurate, and reliable prediction of scramjet flowfields.

Heat addition with variable area: Methodology for preliminary design of the scramjet combustion chamber

Well-known analytical approaches are commonly adopted for the preliminary design of scramjet engines. In this context, the combustion process in the combustion chamber can be modeled by considering heat addition to the airflow at supersonic speed. The one-dimensional Rayleigh flow theory can be applied to estimate the behavior of thermodynamic properties and velocities when the combustion chamber has a constant cross-sectional area and no mass is added within the duct. However, the temperature and pressure predicted by using constant area combustion chambers are too high, implying the necessity of modifications in the cross-sectional area of the chamber to avoid thermal choking and excessive pressure gradients. In this case, the unidimensional Rayleigh theory does not fit anymore. This work proposes an analytical methodology to estimate the airflow thermodynamic properties and velocities for scramjet combustion chambers with cross sections of variable areas by using an iterative algorithm that employs the Rayleigh flow area ratio theory. The analytical results were compared with the two-dimensional computational fluid dynamics analysis using the Reynolds-averaged Navier–Stokes method for both inviscid and viscous flow and considering turbulence effects. The proposed analytical model to estimate the flow behavior in the scramjet combustion chamber predicted results in agreement with the physics of the problem and with the results obtained via numerical simulation. The analytical model cannot predict oscillations in the flow properties caused by the expansion waves and their reflections. Still, the behavior and intensity of the properties are well captured along the entire length of three combustion chambers with variable area. The proposed algorithm is also applied to determine the angle of the combustion chamber that allows guaranteeing a constant, or a quasi-constant, static pressure along the length of the combustion chamber, approaching better the ideal thermodynamic Brayton cycle. The proposed model is suitable for preliminary scramjet designs and can be used to solve other problems involving variable area ducts.

Effect of methanol and ethanol on n-decane pyrolysis process by ReaxFF simulation and experiment validation

Alcohol assisted hydrocarbon fuels regenerative cooling is a promising method to deal with scramjet engine heat management of hypersonic flight vehicle. To reveal the effect of alcohol on hydrocarbon fuels pyrolysis pathway, ReaxFF molecular dynamics simulation was done to investigate the variation of intermediate fragments in methanol/ethanol assisted n-decane pyrolysis process. Results shows that, two alcohols increase the initial cracking temperature of n-decane. Two alcohols turn the first-step cracking products of n-decane from C3H7 + C7H15 to C5H11 + C5H11, and make n-decane cracking into C2H5 + C8H17 possible. Initial pyrolysis of n-decane with methanol and ethanol can be coupled by CH3 and C2H5 free radical respectively. In the subsequent pyrolysis process, C10 are completely consumed within 10 ps. C7-C9 fragments are transient intermediates and can hardly accumulate. Experiment indicates that 450–500 °C is the decisive temperature range for n-decane cracking, and adding methanol or ethanol increase the initial cracking temperature of n-decane pyrolysis.

Investigation on the Flow Instability of Supercritical Hydrocarbon Fuels in Cooling Channels

Flow instability in regenerative cooling channels is an important issue for the thermal protection of hypersonic scramjet engines. Taking into account the dynamic process of the heat transfer and flow instability, a one-dimensional transient model with several modules (including the cracking reaction, convective heat transfer, and rapid calculation of thermal properties) has been developed to investigate the flow instability characteristics of supercritical hydrocarbon fuels in cooling channels. The calculated results were compared and validated against the available experiments and numerical benchmarks, attaining good agreements. By virtue of the transient simulations, the dynamic flow patterns under different flow rates were studied in a single cooling channel with -decane being the working substance. Then, the influences of the operating pressure and heated length on the in-tube flow were further investigated. In addition to the Ledinegg instability, several dynamic instability modes were detected under different external driving forces. It was also observed that under a specific range of pressure drop, the in-tube flow could transition from the density-wave oscillation to a new steady state. Moreover, this flow excursion was more likely to be triggered when decreasing the operating pressure or channel length.

Effect of mainstream oxygen concentration on supersonic film cooling and friction reduction performance using hydrocarbon fuel

One of the effective methods to simultaneously achieve the thermal protection and friction reduction requirement of a scramjet engine is the usage of fuel film cooling. As the combustion process inside the combustor proceeds, the oxygen concentration in the burnt gas gradually decreases and will influence the thermal and friction characteristics of the hydrocarbon film. Therefore, the effects of oxygen concentration in the mainstream on the hydrocarbon film are evaluated by numerical simulations based on RANS framework. The results indicate that the cooling performance of the hydrocarbon film is not sensitive to the oxygen concentration, although the fuel film burns more intense as the oxygen concentration increases. While the friction reduction performance is sensitive to the oxygen concentration. Specifically, about 10% friction reduction can be obtained for 10% oxygen in the mainstream compared to that of inert cases, and the wall friction can be further reduced by 30% for 20% oxygen. Further analysis reveals that the mechanism of the above two drag reductions is different. For weak boundary layer combustion, the decrease of the molecular viscosity contributes to the reduction of friction. While for significant boundary layer combustion, a lower near-wall velocity gradient is the main reason.

On tert-butyl hydroperoxide in ignition intensification of n-heptane under engine-relevant condition

The higher reactivity of hydrocarbon fuels is necessary to match supersonic airflow in scramjet engines and micro-scale effect in micro-internal combustion engines. Tert-butyl hydroperoxide (TBHP) was selected in this study as an additive reactively enhancer to explore its capacity to intensify the auto-ignition of a typical hydrocarbon (n-heptane). Ignition delay times (IDTs) of n-heptane doped with TBHP were thus measured at engine-relevant conditions covering pressure of 20 atm, temperatures of 650 – 1250 K, and equivalence ratios of 0.5 and 1.0. A nonlinear promotion effect of TBHP on the auto-ignition of n-heptane was observed. Detonation caused by TBHP in region 2 has an impact on actual reflected shock temperature and pressure at high beginning temperatures (> 1000 K). With the addition of TBHP, the negative temperature coefficient (NTC) behavior of first-stage ignition eventually vanishes accompanied with a weakened main NTC of IDT. Analyses of exothermic reactions, reaction pathways, and species evolution indicated that the quick production of ȮH and ĊH3 radicals from TBHP enhances the reactivity of n-heptane and affects the main exothermic reaction in the early period. This study reveals the potential of TBHP to be used in the ignition regulation of the combustion of hydrocarbon fuels and provides a methodological approach for the efficient utilization of hydrocarbon fuel in scramjet engines and micro internal combustion engines.

Effects of combustion on the near-wall turbulence and performance for supersonic hydrogen film cooling using large eddy simulation

Supersonic film cooling using fuel on board is an effective way to simultaneously shield the huge heat and momentum flux transported from the mainstream to the wall in a scramjet engine. The self-ignition and combustion of the injected fuel film will significantly change the turbulent transport behavior in the boundary layer. To reveal the effects of the boundary layer combustion on the near-wall turbulence and wall fluxes, large eddy simulations (LES) of the Burrows–Kurkov supersonic combustion experiment using hydrogen as a film are performed based on the in-house solver scramjetFoam. The solver successfully captures the additional skin friction reduction phenomenon induced by the boundary layer combustion compared to other numerical works using LES in the public literature. The results reveal that further increased anisotropy of turbulence combined with the low-density region contributes to a remarkable suppression of turbulent transport processes in the wall-normal direction. The self-ignition point of the hydrogen film is found to oscillate back and forth in a span of 80 mm, which significantly enhances turbulence in the boundary layer. However, the increased turbulent fluctuating velocity is mainly concentrated in the streamwise direction, while the other two components are suppressed instead. The findings are also essential for improving engineering computations based on the Reynolds averaged simulation method.

Trajectory Optimization of a Partially Reusable Rocket–Scramjet–Rocket Launch System Including Fly-Back

The integration of high-speed airbreathing engines into dedicated small satellite launchers may allow for some of the stages of the launch system to be reused, resulting in improved cost efficiency and turnaround times. Designing a trajectory of this type is challenging, due to the high number of competing performance drivers and design constraints, which are often in opposition with each other between the different modes of flight (i.e. airbreathing, rocket, and glide). In this work the design of the trajectory for a partially airbreathing rocket–scramjet–rocket multistage launch system is studied by generating optimal trajectories using the pseudospectral method of optimal control for maximum payload-to-orbit. This optimization simultaneously determines the ascent of all three stages, as well as the return trajectory of the scramjet stage within the trajectory optimization. It is found that the scramjet accelerator is able to achieve fly-back to its initial launch site, but that it must ignite its scramjet engines during its return flight, causing this to become an important consideration in the optimal trajectory design. This work shows that the return of the scramjet stage of a rocket–scramjet–rocket launch system may be feasible, and indicates the trajectory features that may heavily influence future partially airbreathing launch system designs.

Characteristics of self-excited shock wave oscillations in a supersonic isolator with exit duct configurations

The structure and dynamics of multiple shock waves inside a constant area isolator of a model scramjet engine are studied numerically using a finite volume solver. The flow is accelerated through a nozzle with an exit Mach number of 1.75 and continues in the constant area isolator where multiple shock waves are formed. To analyze the effect of duct acoustics on the dynamics of shock train oscillations, three ducts with a duct length of 0D, 50D, and 150D are attached to the exit of the isolator, where D is the isolator exit height. The wall static pressure variation along the bottom wall and four specific locations are monitored. The wall shear stress variation along the wall provides the location of separation and reattachment points in the flow field. The statistical analysis provided more insights into the oscillation mechanisms. The lower frequency components from the duct resonance are enhanced as the duct length is increased and it altered the dominant frequency of shock oscillation. The dependence of shock train length on the exit duct confirmation is analyzed by the pressure ratio method. The different frequency components in the flow field are analyzed using the power spectral density of wall static pressure fluctuations and are distinguished based on the frequency range. The cross-spectral analysis of the wall pressure fluctuations provided the correlation between pressure signals and their narrowband time delay. Thus, the origin and directions of the perturbations are obtained and discussed. The increase in exit duct length decreases the downstream disturbance velocity, indicating the duct resonance effect on shock train oscillations. • Simulation of the supersonic isolator flow field to understand the duct resonance effect on self-excited oscillations of shock train. • The duct resonance is varied by attaching ducts of different lengths to the exit of the isolator. • As the duct length is increased, the amplitude of shock oscillation is increased and the dominant frequency of shock oscillation is decreased. • The length of the shock train is increased with an increase in duct length. • The cross-spectral analysis of the wall pressure fluctuations provided the correlation between pressure signals and their narrowband time delay.

Grid Adaptive Technique for Simulation of Scramjet Intake-Isolator at Hypersonic Speeds

Hypersonic intake is one of the major components of Scramjet engine. It compresses the incoming hypersonic flow through a series of oblique shocks as the flow passes through intake-isolator section before entering the combustion chamber, which is essential for efficient combustion. The shocks generated inside in the intake interacts with boundary layer following shock boundary layer interaction and flow separation. The separated flow blocks the flow capture area such that engine expresses unstarting phenomenon. Understanding and mitigating such flow phenomenon is a challenging task. With respect to hypersonic speeds the experimental facilities are very limited. The only alternative to solve this problem is Computational Fluid Dynamics because of its capabilities. But validation of CFD results with analytical or experimental is the foremost prerequisite to chase computational analysis. Mostly at high speeds the precision of CFD results rest on the type of grid, number of elements and turbulence model used. So, in this paper, computational analysis of hypersonic intake is carried out through designed conditions to ensure the correct CFD process is used by varying number of elements in fluid domain by grid adaptive technique using ANSYS Fluent and satisfying Y+ parameter. The domain is analysed with various turbulence models and among them SST has predicted all the flow characteristics of scramjet intake-isolator at hypersonic speeds like separation bubble, shock reattachment, cowl shock etc similar to experimental results with the help of grid adaptive technique. So, grid adaptive technique is also proposed for simulation of scramjet intake at off-design conditions.

Progress of Coupled Heat Transfer Mechanisms of Regenerative Cooling System in a Scramjet

The feasibility of regenerative cooling technology in scramjet engines has been verified, while the heat transfer behavior involved in the process needs further study. This paper expounds on the necessity of coupled heat-transfer analysis and summarizes its research progress. The results show that the effect of pyrolysis on heat transfer in the cooling channel depends on the heat flux and coking rate, and the coupling relationship between combustion and heat transfer is closely related to the fuel flow rate. Therefore, we confirm that regulating the cooling channel layout according to the real heat-flux distribution, suppressing coking, and accurately controlling the fuel flow rate can contribute to accomplishing the optimal collaborative design of cooling performance and combustion performance. Finally, a conjugate thermal analysis model can be used to evaluate the performance of various thermal protection systems.

Effects of wall thermal state on shock wave/boundary layer interaction with foreign-gas supersonic film injection

Shock wave/boundary layer interaction (SWBLI) is considered to be one of the most complex physical phenomena in supersonic flowfield. The flow characteristics of SWBLI can be influenced by wall thermal state. However, in scramjet combustors, with supersonic film injected into the boundary layer, the effects of wall thermal state on SWBLI flow characteristics may have different characters. In this paper, the effects of wall thermal state on SWBLI with foreign-gas supersonic film injection are numerically investigated. During the numerical simulation, considering the actual application of scramjet engines, hydrogen and high-molecule gaseous hydrocarbon fueled supersonic film injections are taken into consideration. The results indicate that, either with or without supersonic film injection, wall thermal state changes both the wave structure and the flow characteristics. Furthermore, the interaction and separation length decrease with decreasing wall cooling heat flux. It is found that, the effects of wall thermal state on the interaction length are determined by its effects on the subsonic layer thickness. Moreover, decreased wall heat flux will cause momentum transport enhancement, which increases both near-wall momentum and wall shear stress, and consequently decreases the separation length. The interesting thing is that, foreign-gas injections lead to different relative variations of the interaction length and the separation length with wall thermal state. The effects of foreign-gas injections on the relative variations of the interaction length under different wall thermal states essentially depend on the foreign-gas specific heat capacity. Meanwhile, the relative variation of the separation length depends on the effects of foreign-gas injection on the relative density variation. It is worth mentioning that, due to the temperature range decrease caused by film injection, low-molecular hydrogen injection has significantly larger separation length relative variations compared with the situation without film injection and with hydrocarbon fuel injection. However, for high-molecular foreign-gas, the relative density variation depends on its own density change with temperature due to its high-density property.

Fluid and combustion dynamics in dual-mode scramjets

Burgeoning technological advancements in practical and efficient hypersonic flight is intertwined with the research and development of airbreathing hypersonic propulsion, specifically dual-mode scramjet (DMS) engines. Due fundamentally to the lack of complete understanding and adequate modeling of the fluid dynamics and combustion processes present in DMSs, a large volume of academic works has been established towards characterizing the physical phenomena present in these engines. Significant differences in flame topologies, fluid interactions, and pressure profiles between scram and ram combustion are observed across these experimental and computational works. A focus on the dynamics responsible for combustion mode transition, choking and the propagation of the pseudoshock, is made, as is a discussion on the theoretical underpinning of the mechanisms behind flow choking and important choking thresholds. Further insight into the fundamental mechanisms and fluid and combustion physics present in DMSs may improve future designs and operability of dual-mode scramjet engines.

Approximate Thermal Analysis of Heating and Deformation of a C/C-SiC Plate in a Scramjet Engine

Approximate Thermal Analysis of Heating and Deformation of a C/C-SiC Plate in a Scramjet Engine

Numerical modelling of braided ceramic fiber seals by using element differential method

• A multi-physical coupled model is proposed for the braided ceramic fiber seals. • The coupled fields contain the process of heat conduction, seepage and deformation. • Element differential method (EDM) is used to solve the multi-physical problems. • The material-dependent parameters in model are obtained by an inverse procedure. • Compared to FEM, EDM is more easy and efficient dealing with the coupled model. A new kind of dynamic seal with braided ceramic fibers has been designed to seal the movable panels in the scramjet engines, which are used for the propulsion of hypersonic vehicle. The braided ceramic fiber structure can provide buffer forces when the seals are subjected to the external dynamical preloads. However, it also makes the seals difficult to analyze. Up to now, the analysis of the seals still uses 1D models so that one cannot implement the coupled analysis of seals and their surrounding structures and flows. In this paper, a novel 2D and 3D mechanics-thermal-seepage coupled model is proposed to describe these seals, which provides a possibility for the aforementioned coupled analysis. Meanwhile, a strong-form numerical method, element differential method (EDM) is employed to discretize the governing equations of the coupled model due to its efficiency and robustness. Three examples are given. The first one is to invert the material-dependent parameters of three kinds of seal strips from the experimental data by Levenberg-Marquardt (LM) algorithm. Other two implement the analyses of 3D square and circular cross-section seals, respectively, which verify that EDM is more efficient than FEM in seal analysis when using the same mesh sizes.

Experimental Investigation on Gliding Arc Plasma Ignition and Assisted Combustion Actuator

In a scramjet engine, the viscosity of kerosene increases, the oxygen content decreases, the atomization evaporation rate and chemical reaction rate of kerosene decrease significantly, and the flame propagation speed decreases significantly under the condition of low total temperature, leading to the difficulty of ignition. Therefore, it is urgent to innovate the repeatable ignition method suitable for ramjet from the source. Gliding arc plasma has the advantages of both thermal equilibrium and nonequilibrium plasma. It can be used as a stable high-temperature heat source and chemical reaction source to strengthen the whole process from ignition to flame combustion, which has a critical application prospect in broadening the ignition boundary of scramjet engines at low total temperature and improving flame stability. In this article, a gliding arc plasma igniter is designed and developed. The research of gliding arc plasma ignition technology in scramjet combustor is carried out from two aspects of the discharge characteristics of gliding arc plasma igniter and kerosene cracking characteristics experiment. The results show that the average power of the gliding arc igniter in the air environment is higher than that in the nitrogen environment. With the increase of flow rate, the average power of the igniter increases first and then decreases. The maximum average power of the igniter with nitrogen is 193.56 W, and the maximum average power in the air environment is 323.49 W. The cracking products of kerosene are mainly hydrogen, methane, ethylene, acetylene, and other hydrocarbons below C3. The concentration of the product was positively correlated with the excitation voltage and channel length, and the concentration of hydrogen was the largest. The concentration of cracking products is higher in air, where hydrogen is three times higher than that in nitrogen. The main flow ignition of scramjet was successfully achieved under the condition of 2-Ma inlet velocity and 1000 K inlet temperature.

Research on Transient Characteristics of Scramjet Engine during Starting Sequence

Transient characteristics of entire engine system affect control method of the engine system. A quasi-transient simulator called Hypersonic Engine Dynamic Simulator (HEDS) was developed for the purpose of analyzing the transient behavior of a scramjet engine system. The simulator was constructed using the volume junction method, which is a type of finite volume method, and was introduced with a parallel computing function using the MPI library. In this study, the transient characteristics at the starting sequence of a scramjet engine equipped with an expander cycle was analyzed. In the startup transient analysis, two factors (the time from high-speed airflow inflow to the main valve opening and the time from the main valve opening to ignition) that affect the system startup were focused. In addition, constraints on the system were set to evaluate results of the analysis. As a result, there existed cases in which the maximum rotational speed of the turbo pump was much higher than expected when the interval between the inflow of high-speed airflow and the opening of the main valve was increased. In addition, there was also a delay time between ignition and fuel temperature rise. Therefore, the time from the main valve opening to ignition should be as short as possible. In the control phase, it was necessary to manipulate those two factors appropriately in order to bring the system to a steady state in a short time. HEDS was found to be a very effective tool for understanding the behavior of the system at the operating point and constructing a control method.