Scramjet Engine Research Articles

Effects of Oxygen Enrichment on Supersonic Combustion in a Mach 10 Scramjet

Scramjet engines are a suitable replacement for rockets in access-to-space systems, but, being air-breathing engines, their operational envelope is limited to altitudes where enough oxygen is available for combustion. Oxygen enrichment, in which fuel is premixed with oxygen, is a promising technique to overcome constraints in altitude due to low oxygen availability, and the difficulty to generate positive thrust at high Mach numbers. Large-eddy simulations of an inlet-fueled Mach 10 scramjet engine are performed to investigate the effect of oxygen enrichment on the flowfield and to characterize the turbulent combustion regimes. It is shown that combustion levels increase substantially with the addition of premixed oxygen, which leads to a considerable increase in combustion efficiency. It is shown that the mixing process is altered by oxygen enrichment in such a way that the observed increase in mixing efficiency is higher than the increase caused by premixing alone. Finally, the combustion regimes in the simulated engine are analyzed. This shows that combustion is mostly nonpremixed, and oxygen enrichment leads to an increase in premixed combustion, but without significantly altering turbulent combustion overall, which is distributed over a large range of regimes.

Control of the oscillations of shock train using boundary layer suction

The isolator plays a major role in the scramjet engine by providing suitable pressure for the combustion process. It is situated between the inlet and the combustion chamber. The deceleration of supersonic flow into subsonic flow is achieved by a series of bifurcated normal shock waves referred to as “shock train”. The flow field is more complex with shock–shock interaction and shock boundary layer interaction. The presence of such flow inside the isolator can degrade the performance of the scramjet engine. The present study focuses on the self-excited shock train flow in an isolator and its control by partial removal of the boundary layer. The results examine the variation of the inlet to outlet pressure ratio with and without suction flow control. Numerical results indicate that the boundary layer suction will cause downstream movement of the shock train and the length of the shock train is reduced. The transformation of the shock train into a single curved normal shock is observed when the shock train present closer or upstream of the suction slot region. The obtained result shows the self-excited oscillation frequency in the range of 14-15 Hz. The suction control provides an advantage by 50% reduction in the amplitude of the shock train movement.

Multipoint Design Optimization of Busemann-Based Intakes for Scramjet-Powered Ascent Flight

Scramjet engines are one of the most promising hypersonic airbreathing propulsion technologies for robust, efficient, and economical access to space. Multi-objective design optimization has been conducted for Busemann-based intakes in inviscid and viscous regimes at single and multiple design points, respectively, by means of surrogate-assisted evolutionary algorithms coupled with computational fluid dynamics in the present research. Intake geometries are generated by applying geometric alterations to the full Busemann intake via leading-edge truncation, axial stunting, and radial contraction, aiming to simultaneously minimize intake drag and maximize the compression efficiency at two different design conditions (i.e., Mach 7.7 at an altitude of 30 km and Mach 10 at 33.5 km on a constant dynamic pressure ascent trajectory). The single-point inviscid optimization study has found that the nondominated solutions obtained from minimizing drag and maximizing compression efficiency are approximately the same as those obtained from maximizing total pressure recovery and minimizing static pressure ratio. From the multipoint viscous optimization study, the optimal solutions have been found to retain the original advantages of the inviscid full Busemann intakes in terms of high compression efficiency with shorter intakes and higher static pressure as well as adequately high mean exit temperature for both design conditions.

Hydrogen fueled scramjet combustor with a wavy-wall double strut fuel injector

The demand for the enhancement of scramjet performance is increasing worldwide due to its tremendous and upcoming applications. The major challenge in scramjet engines is the mixing of fuel and air and succeeding combustion at supersonic speed. In a scramjet combustor flow field, the development of high turbulence (flow disturbance), shear-mixing layer, and oblique shock waves are the key factors to magnify the mixing efficiency of fuel and air at the significantly less resident time of air in the combustion chamber. Based on these techniques, an innovative strut fuel injector with a wavy wall surface and double strut has been investigated for possible mixing enhancement of fuel and air. These results are compared with the basic single wedge strut available in the open literature to identify the performance improvement. The computational geometries are modeled with Reynolds-averaged Navier Stokes equations, SST k-ω model, and global one-step reaction mechanism. All the flow modeling has been solved with finite volume and a second ordered upwind discretization scheme. The evaluation of scramjet internal flow structure and numerical results found that the double strut with a wavy-wall shape improved the mixing efficiency and combustion phenomenon significantly. The mixing and combustion profiles reveal that wavy-wall double strut generates an early complete mixing within a distance of 0.17 m with increased mixing and combustion efficiency of 18% and 20%, respectively.

Effects of fuel decomposition and stratification on the forced ignition of a static flammable mixture

In advanced propulsion systems such as scramjet engines, endothermic decomposition of onboard large hydrocarbon fuels can be used effectively for cooling and active thermal protection. During the cooling process, large hydrocarbon fuels absorb heat and decompose into small fragments. Since fuel decomposition changes the chemical and transport properties of the reactants, it is expected to affect the combustion afterwards. In this study, forced ignition in quiescent n-decane/air mixtures with fuel decomposition is investigated via a simplified model and transient numerical simulations considering detailed chemistry and transport. The emphasis is placed on assessing the effects of fuel decomposition on ignition kernel development and minimum ignition energy (MIE) in both homogenous and fuel-stratified mixtures. Fuel decomposition is modelled by a homogeneous ignition process in n-decane/air mixture at constant atmospheric pressure and with an initial temperature of 1300 K. Small fragments appear during the pyrolysis process. The partially-reacted mixture is frozen and cooled to a lower temperature and used as the initial mixture in forced ignition. For homogeneous mixtures, fuel decomposition can greatly promote ignition for fuel-lean decane/air mixtures while it has little effect for the stoichiometric case. Fuel decomposition also affects the duration of unsteady ignition kernel transition. Besides, fuel decomposition and fuel stratification are combined to further promote forced ignition. New flame regimes are observed and an optimum stratification radius is identified. In order to promote the forced ignition, only the fuel within the optimum stratification radius needs to be decomposed. Furthermore, laser and spark ignition experiments are conducted to measure the MIE of n-decane/ethylene/air mixtures with different equivalence ratios and ethylene blending ratios. The MIE measured in experiments cannot be directly compared with simulation results since the simulation model for forced-ignition is simplified. Nevertheless, the experimental results are consistent with simulation results and thus validate some conclusions mentioned above. The present results provide useful guidance to the fundamental understanding of forced ignition in a mixture with fuel decomposition.

Experimental validation of a new type of direct-connect facility for scramjet studies

Combustion chamber design has been important to the development of scramjet engines for a long time. Conventional experimental facilities of scramjet could be divided into two types: free-jet and direct-connect facilities. Free-jet facility can provide real incoming conditions to the combustion chamber, especially the real shock wave train structures formed in inlet and isolator. Unfortunately, it is relatively expensive. While direct-connect facility has a distinct advantage in experimental cost and period, so it has been used widely in relevant researches. However, traditional direct-connect facility usually can only provide a uniform incoming flow, which is much different from the distorted flow field in real scramjet, and this is a non-negligible disadvantage to relevant design. An inverse design method based on method of characteristics (MOC) is presented to solve this problem. The facility designed by this new method can produce the pre-assigned distorted flow field without encountering unstart phenomenon. In this paper, experimental investigation of the new direct-connect facility, which is designed to reproduce a specific distorted flow field with Mach number more than 2, is further presented. Test data of particle image velocimetry (PIV), flow visualization experiments and corresponding numerical simulation results are used to verify the reliability and accuracy of this new facility. Qualitative flow visualization and quantitative measurement results further prove that this new type of direct-connect facility can generate the target distorted flow field as pre-defined.

Output tracking for nonlinear systems subject to unmodeled sluggish actuator dynamics via model-based extended state observer

We propose an output tracking control approach for a class of uncertain, nonlinear systems with unmodeled, sluggish actuator dynamics. The normal extended state observer (ESO) is modified with a nominal model of the largely unknown actuator dynamics. The total disturbance is redefined to include the uncertainties of the actuator dynamics. The feedback law is correspondingly designed to cancel the estimated total disturbance. The closed-loop system is guaranteed with bounded stability through theoretical analysis. The proposed controller is extensively validated and compared to normal active disturbance rejection control (ADRC), nonlinear dynamic inversion control (NDIC), and L1 adaptive control by a numerical example, and a practical application to both the attitude control and velocity control of a hypersonic vehicle, where the deflection of the elevator and the thrust variation of the scramjet engine are considered.

Modeling of the non-uniform combustion in a scramjet engine

Due to the high-speed of the air stream, the scramjet combustion is neither uniform nor complete. An empirical fuel distribution model is developed to describe non-uniform combustion for scramjet engines. The combustor is subdivided into different regions by radius with different local equivalence ratios. The conservation equations of the regions for the combustion and exhaust expansion are computed independently. The results indicate that scramjet thrust is more related to the fuel equivalence ratio and combustion efficiency. If the combustion efficiency is 100% and the fuel equivalence ratio is constant, there is no obvious effect of different fuel distribution to the engine except for the exhaust parameter distribution. It is also revealed that the sum of isolator shock loss and combustor Rayleigh loss is nearly constant under the same isolate cross-sectional area. Lower isolator inlet Mach is benefit to the thrust performance and the best thrust performance is at the thermal choking boundary. When the isolator inlet Mach increases, the thrust decreases.

Inhomogeneous behavior of supercritical hydrocarbon fuel flow in a regenerative cooling channel for a scramjet engine

Supercritical hydrocarbon fuel is employed to simultaneously cool associated combustors and to recover heat. Aiming to improve the thermal performance of the regenerative cooling channels in an advanced aircraft, the flow behavior, vortex structures, and transport characteristics of supercritical n-decane are investigated using an LES model. The flow patterns are identified by several common buoyancy criteria, and oscillation effects are observed in strongly mixed flows. Using temperature and density gradient fields, distinct inhomogeneity shows that the thermal oscillations aggravate the further occurrence of uneven performance. Furthermore, the positive helicity gradually increases with the accumulation of buoyancy force. The notably strong vorticity mainly concentrates near the heated wall, while a narrow section appears in the turbulent region, which indicates that the variable vorticity behavior contributes to the transport properties.

Synergistic promotion of multi-component RP-3 kerosene low-temperature ignition with addition of radical activator

Scramjet engines play a vital role in hypersonic vehicles, but they suffer from the ignition issue due to the low reactivity of hydrocarbon fuel when flying at low flight Mach numbers. In this work, the low-temperature ignition performance of multi-component kerosene was significantly improved by only adding 3% borane-dimethyl sulfide (BDMS), resulting in the low-temperature ignition reducing greatly from 620 to 160 °C. This indicates that BDMS was an effective activator for multi-component kerosene. To analyze the low-temperature efficient activation mechanism of multi-component kerosene, paraxylene, ethylcyclohexane, n-decane, and n-tetradecane (each with added 3% BDMS) were studied as they are essential components in kerosene, respectively. The low-temperature ignition performance and the ignition process of n-alkane were significantly improved and changed for a single component of hydrocarbon. There was a synergistic promotion effect for multi-component hydrocarbons with the addition of activator BDMS. BDMS could significantly improve the low-temperature ignition performance and oxidation of ethylcychohexane in the presence of n-tetradecane. The addition of paraxylene promoted the evaporation of the hybrid fuel of n-tetradecane/ethylcyclohexane. The results obtained from this study could help to gain a better understanding of the low-temperature ignition of multi-component kerosene and provide an effective method to solve the scramjet ignition problem at low flight Mach numbers.

Estimation of Aerodynamic Heating on Scramjet Inlets and Validation With Measurements

Abstract Aerodynamic heating levels on a typical inlet configuration of a scramjet engine are estimated using both standard design correlations and numerical flow simulations. The stagnation point heat flux is estimated using the Fay and Riddell formula. Aerodynamic heating over the inclined ramps is estimated using Van Driest method. Numerical flow simulations are carried out using a Reynolds averaged Navier–Stokes (RANS) solver coupled with energy equations and the Shear Stress Transport (SST) k–ω turbulence model. The aerodynamic heat flux estimates are validated with in-house measurements in a shock tunnel and for a scramjet flight experiment in the Mach number range 1.59 to 7.92. The emergence of a good agreement between them confirms the appropriateness of design correlations for heat flux estimation in scramjet inlets. The choice of simplification and appropriateness of design correlations to complex geometries demand critical assessment. Numerical flow simulations capture flow features and enable the identification of potential augmented heating zones, which will be critical for long-duration scramjet missions.

The spatial growth of supersonic reacting mixing layers: Effects of combustion mode

Supersonic mixing layer flows are taken in scramjet engines for enhancing mixing and combustion efficiency. The spatial growth of supersonic reacting mixing layers with streams of dilute hydrogen and air has been investigated using direct numerical simulation, with detailed chemical reaction mechanisms. Different inlet air temperatures are considered to trigger auto-ignition of mixtures in mixing layers with Damköhler numbers ranged. By the comparison of auto-ignition sites and vortex shedding positions in the computational domain, four modes are classified as incombustible, premixed combustion, diffusion combustion after premixed one, and diffusion combustion. The spatial growth of supersonic reacting mixing layers, characterized by the vorticity thickness, along the streamwise direction establishes a strong relationship with the combustion mode. For the premixed combustion mode the growth rate increases compared to inert case, but it decreases for the diffusion combustion mode. The difference of combustion mode implies that the spatial growth has to be viewed section by section rather than previous one as a whole. The premixed combustion with strong heat-release within the shear layer causes much dilatation, while the diffusion combustion with relatively weak heat-release suppresses entrainment of the shear interface to an extent so that offsets dilatation influences. These findings provide a new perspective for supersonic reacting mixing layers applied to scramjet engines.

Mixing and transient combustion processes of scramjet combustor with transverse injector and hypermixer

The supersonic mixing between the reactive injection with an airstream is an essential issue that determines the scramjet engine performance. The mixing and transient combustion processes under different equivalence ratios for both the transverse injector and hypermixer in a benchmarked Hyshot combustor are numerically investigated with a Mach 2.49 inlet flow. The improved delayed detached eddy simulation (IDDES) coupled with the Thickened Flame Model and seven species and seven reactions for the H2/air reaction mechanism was used to investigate differences in the mixing, shock structure, combustion mode, and combustion efficiency between the transverse injector and hypermixer. The results show that interactions between the streamwise vortex and hypermixer significantly impact the mixing and combustion enhancements. The combustors with the hypermixer and transverse injector are both in the scramjet mode at the equivalence ratio of 0.295. As the equivalence ratio increases to 0.593, the combustor with the hypermixer is still in the scramjet mode, but the combustor with transverse injector is already in the ramjet mode. The numerical results show that compared with the general transverse injector scheme, the scramjet combustor with the hypermixer has higher mixing and combustion efficiencies.

Heat transfer deterioration in upward and downward pipe flows of supercritical n-decane for actively regenerative cooling

In this paper, we consider the flow and heat transfer behaviour of turbulent upward and downward flows of supercritical n-decane, in order to reveal the features of heat transfer deterioration (HTD) that would be expected in relevant active regenerative cooling systems for scramjet engines. Specific focus is placed on key velocity-field features that appear in these flows. Following the validation of six turbulence models, the SST k-ω and RNG k-ϵ models are found to be suitable for simulating the upward and downward flow cases, respectively. “M” type velocity profiles (a non-monotonicity of the velocity along the radial direction) are observed, which arise due to a spatially-varying interplay between the inertial and viscous forces in the flow domain, while larger velocity gradients in the buffer layer are also observed that contribute to the phenomenon of HTD. Furthermore, it is found that the secondary flows as well as the different mass fluxes that arise due to the velocity increase from the wall to the flow core zone (i.e., the influencing range and intensity of cross-sectional kinetic energy), respectively, are observed in the HTD development region, as well as the HTD peak area and degradation regions. A zone of higher thermal diffusion appears in the near-wall region, which acts as a thermal barrier and contributes to HTD.

Heat transfer characteristics of kerosene phase change based on fuel flow rates in the regenerative cooling heat exchanger of scramjet engines

Heat transfer characteristics of kerosene phase change based on fuel flow rates in the regenerative cooling heat exchanger of scramjet engines

Reduction of surface friction drag in scramjet engine by boundary layer combustion

This study investigates the effects of boundary layer combustion on skin friction reduction in a model scramjet combustor. The 4-equation RANS model (Transition SST model) is employed as the turbulence model and one experimental case which involves supersonic turbulent boundary layer combustion is used for validating the numerical simulation method. Then a wall jet device which can be used to inject hydrogen into boundary layer is designed and added to the lower wall of the combustor in the model scramjet engine. The numerical results showed that the optimal ratio between the primary fuel, producing thrust, and the wall jet fuel, used for drag reduction, for getting the largest skin-friction drag reduction when installing the wall jet device in the combustor, is 3:1. The study of modifying the location of the wall jet device showed that setting the wall jet position too close to the upstream or downstream could not get the maximum overall drag reduction. The installation position in the flow path depends on the coupling of primary fuel and wall jet fuel combustion. In practical scramjet design, the distribution ratio of primary to wall jet fuel and the location of wall jet device should be weighed to improve skin-friction reduction efficiency and overall engine performance.

Performance characteristics of a scramjet engine using JP-10 fuel containing aluminum nanoparticles

High energetic nanoparticle as additives is of great significance for liquid hydrocarbon fuels that will exhibit high heat value and potential performance improvement in hypersonic and related combined circle propulsion system. In the study, the JP-10 slurry fuel containing 16 wt% aluminum nanoparticles was experimented upon a scramjet engine to investigate the combustion efficiency and engine performance in comparison to pure JP-10 fuel. It is the first to realize the stable combustion of hydrocarbon fuel containing nanoparticles in scramjet engine. The wall pressure profile and calculated Mach number along the combustor are firstly compared and analyzed with the equivalence ratio of 0.56, 0.73 and 0.91. At the equivalence ratio of 0.56, the addition of aluminum nanoparticles promoted the engine to transition from the scramjet mode to the dual-model ramjet mode in advance, and the wall pressure along the combustor path to increase significantly. The combustion efficiency data of the two fuels are calculated by one-dimensional impulse analysis method. The results show that the combustion efficiency of slurry fuel is significantly higher than that of pure fuel. Besides, the addition of aluminum nanoparticles also enhanced the density specific impulse. When the equivalence ratio is 0.56, 0.53 and 0.91, the density specific impulse of slurry fuel increases by 34.37%, 11.33% and 8.96% respectively. The above research indicates that the use of slurry fuel containing aluminum nanoparticles indeed improves engine performance and has great potential in engineering applications.

Hysteresis of Shock Train Movement in the Isolator with a Ramp

Inside the dual-mode scramjet engine, the shock train will move to a new location when the backpressure changes. Few works focus on the response of the shock train position to the backpressure cyclic variation. This work aims to investigate the behavior of the shock train under such backpressure conditions. Experiments were carried out in a Mach 3 direct-connect facility. The isolator is equipped with a ramp that is used to improve the isolator performance. The static pressures along the wall centerlines were measured. The schlieren imaging was used to provide flow visualization. The results show that a significant hysteresis occurs in the shock train position during the backpressure cyclic variation process. It is found that a large-scale subsonic wake flow region forms behind the ramp when the shock train reaches the ramp trailing edge. The capability of the ramp to retain the existence of the wake flow determines the occurrence of the hysteresis. The effects of the ramp height and width on the hysteresis were examined. Based on the experimental data, the oscillation characteristics were discussed by using wavelet analysis and cross-spectrum analysis. The coupled oscillation between the shock train oscillation and the backpressure oscillation was observed.

Mixing performance of transverse hydrogen/air multi-jet through coaxial injector arrays in supersonic crossflow

Increasing the fuel mixing performance substantially improves the overall performance of the scramjet engine for a long-distance flight. In this paper, the influence of coaxial injector arrays of hydrogen/air multi-jet on the mixing performance of the fuel in supersonic crossflow is fully investigated. Our main goal is to examine the impacts of air and fuel coaxial injector on fuel distribution and penetration downstream of injectors in different operating conditions. In this study, fuel and air are simultaneously injected through coaxial multi-jet at sonic condition while of free-stream Mach number is 4. Computational Fluid Dynamic is applied for simulation of the transverse coaxial jet at supersonic crossflow. The effect of jet diameter with the same mass flow rate of air and hydrogen on fuel mixing is also investigated. The mixing efficiency of different jet spaces and pressures is also examined to obtain an optimum jet arrangement in the combustor chamber. Our study shows that the injection of the coaxial air/hydrogen jet noticeably improves mixing downstream by augmentation of fuel interaction with an air jet. Our results also show that fuel jet space of 7 Dj offers maximum fuel mixing by the formation of multi vortices with uniform strength.

The Effect of Ramp Location in a Strut Based Scramjet Combustor under non reacting flow field

The scramjet engine is considered to be the high-speed propulsive system for hypersonic air-breathing vehicles. Researchers focus on various geometries of strut injectors to enhance the mixing and combustion efficiencies of the scramjet combustor. This paper reveals the non-reacting flow characteristics of various ramp locations in a scramjet combustor with strut injection. The ramps are symmetrically placed at three different locations upstream of the strut within the combustor. Air enters into the combustor inlet with Mach number 2, and the fuel is injected at the sonic velocity from the strut. The 2D supersonic flow characteristics are numerically investigated using ANSYS 18.0 software. The steady-state flow simulations are performed in this study. The flow field is modelled using RANS equations, and the k-ω SST turbulence model and default constants are chosen to investigate the flow characteristics within the combustor. The flow characteristics of the ramp based DLR scramjet combustors are compared with the baseline DLR scramjet model. The computational results are validated with the experimental data. The shock wave interaction from the ramps enhances the distribution of hydrogen in the lateral direction of the flow than the baseline strut configuration. The ramps create recirculation regions that could enhance fuel-air mixing. The ramp induced strut based DLR scramjet engine provides increased total pressure loss of the baseline DLR scramjet.