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Combustion Stabilization, Structure, and Spreading in a Laboratory Dual-Mode Scramjet Combustor.
Abstract and Figures
Dual-mode scramjets have the potential to provide efficient, air-breathing propulsion at high flight Mach numbers. Flame stabilization and spreading are a challenge in such engines due to the very high flow velocities. Combustion occurs in a complex regime where both flame properties and auto-ignition reactions are expected to be important.
The focus of the current study is to improve the physical understanding of the combustion mechanism and its practical implication in such combustors. Topics of interest are the combustion stabilization locations, the detailed structure of the reaction zone, and the physical mechanisms controlling the heat release distribution. A facility was developed for the experimental investigation of a laboratory dual-mode scramjet combustor at conditions equivalent to flight Mach numbers of 4.3 to 5.5. The combustor contained flush wall fuel injection and a cavity flameholder, which are the basic flow elements in many proposed practical designs. The diagnostics used include high speed movies of the chemiluminescence, wall pressure measurements, and planar laser induced fluorescence of CH, and simultaneous OH/formaldehyde.
The study revealed two distinct reaction zone structures that are caused by two flame anchoring locations. Cavity stabilized combustion occurs at low stagnation temperatures. For these conditions the reaction zone is anchored at the cavity leading edge and the flame spreading is controlled by premixed flame propagation. Jet-wake stabilized combustion occurs at high stagnation temperatures. The reaction zone is a lifted jet flame, which has a premixed base and a downstream diffusion flame. In all cases, initial auto-ignition reactions occur well upstream of the primary reaction zone, resulting in an auto-ignition assisted flame base. The results are useful for developing physics based models of the combustion.
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... The recognition of the significant combustion instability characteristics in cavity-based scramjet combustors has drawn ever-increasing attention in the last few decades. To date, it has been found that combustion oscillations in a scramjet could be induced by different mechanisms, including boundary-layer separation [8,9], thermal chocking [10][11][12], autoignition [13,14], and deflagration/detonation transition [15]. ...
... Several questions need to be addressed, including the change rule of combustion oscillations under different conditions (e.g., inflow, fuel injection scheme, and combustor configuration), interactions between different oscillation mechanisms, and the mechanisms for large-scale combustion oscillations such as the flame flashback and corresponding control. Unlike the results reported by Lin et al. [5], Micka and Driscoll [13,19] measured the pressure and flame oscillations in a dual-mode scramjet combustor and did not find the expected sharp peak in the frequency range for instabilities caused by thermoacoustic or periodic shed vortices. Such a broadband power-spectral density distribution reflects a strong nonlinearity [20], which corresponds to the higher complexity of combustion oscillations in scramjets [21]. ...
This work experimentally studied combustion oscillations in a cavity-based combustor operating under high-enthalpy supersonic flow. Three fuel-injection distances and the corresponding combined fuel injection schemes were tested and compared with the global equivalence ratio unchanged. The results indicate that the combustions of longer and medium distances (35 and 21 mm, respectively) present stronger oscillations, which are mainly attributed to the strengthened oscillations of the mainstream flame (MF). It is also found that combined fueling schemes can help weaken the oscillations. The predominant frequencies of the combustion oscillations in different fueling schemes are all distributed at approximately 400 Hz, which are proven to be induced by the MF. The low-frequency range (0-50 Hz) is more dominant for the shorter injection distance (7 mm) than in the other cases, which is because the flames are strongly influenced by low-frequency disturbances with fluctuations in the fueling injection and inflow. The analysis reveals that these disturbances act primarily on the cavity flame. Nomenclature C x;y = magnitude-squared coherence of signals x and y D1 = depth of cavity fore wall, mm D2 = depth of cavity rear wall, mm d = distance between injection location and cavity fore wall, mm d = nondimensional injection distance f = frequency, Hz H = combustor height, mm I = total luminosity intensity of region L = length of cavity, mm Ma = Mach number P = pressure, kPa S = sample standard deviation t = time, ms W = combustor width, mm α = amplitude of dynamic mode decomposition mode Φ = diameter of jet orifice, mm Subscripts CF = cavity flame in = inflow jet = fuel jet MF = mainstream flame mean = averaged result 0 = stagnation condition
This work experimentally studied combustion oscillations in a cavity-based combustor operating under high-enthalpy supersonic flow. Three fuel-injection distances and the corresponding combined fuel injection schemes were tested and compared with the global equivalence ratio unchanged. The results indicate that the combustions of longer and medium distances (35 and 21 mm, respectively) present stronger oscillations, which are mainly attributed to the strengthened oscillations of the mainstream flame (MF). It is also found that combined fueling schemes can help weaken the oscillations. The predominant frequencies of the combustion oscillations in different fueling schemes are all distributed at approximately 400 Hz, which are proven to be induced by the MF. The low-frequency range (0–50 Hz) is more dominant for the shorter injection distance (7 mm) than in the other cases, which is because the flames are strongly influenced by low-frequency disturbances with fluctuations in the fueling injection and inflow. The analysis reveals that these disturbances act primarily on the cavity flame.
To investigate the effect of the multicavity on flame stabilization mode transition in a scramjet combustor, experiments involving various fuel injection strategies were conducted in a direct-connect supersonic combustor with a multicavity at the entrance with a Mach number of 3.0. The flame stabilization mode of the first cavity transitioned under all working conditions, but the equivalence ratio and wall static pressure changes were different. According to the different driving modes, the transition types were identified to be direct-driven local heat release and indirect-driven downstream backpressure. Direct changes in the local equivalence ratio could lead to significant combustion variations. The difference in combustion intensity between the two flame stabilization modes resulted in obvious path dependence, and the different transition paths of the flame stabilization mode could yield a difference in the equivalence ratio of [Formula: see text] at the moment of flame stabilization mode transition. If the equivalence ratio of the first cavity was set to a low value, the backpressure generated by downstream combustion was indirectly employed to promote the flame stabilization mode transition of the first cavity, which could reduce the degree of abrupt change. In addition, different transition paths could generate a difference in the equivalence ratio of [Formula: see text] at the moment of flame stabilization mode transition.
A numerical investigation has been conducted using 2D RANS (Reynolds Averaged Numerical Solution) equation and SST k- ω turbulence model in a cavity-based scramjet combustor with a doubly-dual cavity and hydrogen-fuelled, to study the performance and combustion characteristics of three struts and four struts injection. Fluid flow characteristics of four-strut injectors were compared with that of three-strut injectors. The reason for better performance output from four strut injection is understood and discussed briefly. Compared with the three-strut injector configuration, the formation of recirculation regions and vortices are enhanced in the four-strut injector system affecting shock wave interactions (shock-shear, shock-shock layer). Due to increased levels of shock layer interaction, and lower presence of oxygen, there was a better mixing of the fuel which resulted in improvement in combustion when four strut injectors were introduced. The intensity of combustion also increased which is evident from static temperature plots.
The present study experimentally investigates the transition of flame stabilization mode at Ma = 2.92 using a model scramjet combustor. The combustion mode is analyzed by pressure measurement and quasi-one-dimensional analysis. The effects of both equivalence ratio and aft wall offset on flame behaviors are further characterized by CH* chemiluminescence. The results show that the transition from cavity stabilized combustion mode to jet-wake stabilized combustion mode is realized by increasing the equivalence ratio. And the quasi-one-dimensional analysis reveals that all cases work in the Early Scram mode and approach thermal chocking after increasing equivalence ratio. As the equivalence ratio increases, the amplitudes of the flamebase and heat release gradually increase, and their distributions gradually migrate downstream. The amplitude of flamebase and heat transfer remains prominent with the consistent rise in equivalence ratio. At the same time, the spatial distribution of heat release tends to spread out upstream and occasionally appear upstream of the leading edge of the cavity, indicating that the flame stabilization mode transforms from cavity stabilized combustion mode to jet-wake stabilized combustion mode. The spatial variation of local heat release rate is responsible for the transition of flame stabilization mode. A slight aft wall offset promotes the transition from cavity stabilized combustion mode to jet-wake stabilized combustion mode at a higher equivalence ratio.
Thrust abruption due to mode transition could be catastrophic for hypersonic vehicles. To understand the underlying physics, a direct-connected transient Flight Trajectory Simulator 1 (FTS-1) has been developed at the Institute of Mechanics, Chinese Academy of Sciences. This facility uses advanced high-speed measuring techniques, including thrust and static wall pressure measurement, and Schlieren and hydrocarbon radical chemiluminescence (CH * chemiluminescence) imaging. Kerosene-fueled dual-mode combustor experiments are designed in an acceleration trajectory. The basic operation parameters and the flame and flow dynamics of the acceleration induced mode transition are evaluated. Discussions are given on the triggering mechanisms responsible for the ram to scram mode transition in a simulated flight acceleration.
In this study, the flow dynamics with finite volume approach on commercial software Ansys-Fluent 20.0 to solve the compressible two-dimensional fluid flow with Reynolds Average Navier Stokes equation (RANS) equation by considering the density-based solver with Shaer stress transport model (SST) k- ω turbulent model. The species transport model with volumetric reaction and finite rate/eddy dissipation turbulence chemistry interaction is adopted to study the combustion phenomena. Additionally, the effect of spacing between the struts on the flow characters and performance of the combustor is studied by increasing the spacing of struts from 1 mm to 4 mm for each increment of 1 mm. It is found that the multi strut improves the mixing and combustion efficiency compared with that of the single strut owing to the formation of a significant separation layer, resulting in multiple shocks, vortices, and a larger recirculation zone. However, when the spacing of struts is increased further, the performance of the combustor is found to be deteriorating owing to the formation of larger separation layers. The recirculation zone is significant when the strut spacing is minimal and shrinks and restricts itself within the cavity when spacing is increased. So, for better performance of combustor, multi strut with minimum spacing is preferable.
Supersonic hydrogen-air jet flames were investigated experimentally in order to measure the general flame properties and to investigate the effect of shock waves. This experiment was the first reacting flow experiment interacting with shock waves. The general properties measured were flame length, blowout limits, heat release patterns, static pressure distributions, stagnation pressure losses, thermal choking limits, and nitric oxides levels. An improved Mach 2.5 supersonic combustor was built. The supersonic flames are round to be significantly shorter than corresponding subsonic flames, for the same velocity and density ratios. This difference implies that for axisymmetric jet geometries with combustion, mixing rates are larger in the supersonic case. One possible reason for the relatively short supersonic flames is that the radial velocity, which controls entrainment, can be altered by compression/expansion waves that are inherent in the jet geometry. Shock effects were investigated by changing shock strength and position with particular emphasis on the lengths and stability limits. It was found that shock waves enhance the fuel-air mixing such that flame lengths decreased by 30% when an optimum shock location and shock strength was chosen. Enhanced mixing resulted, in part, because the shocks turn the flow and induce radial inflows of air into the fuel jet. Substantial improvements in the flame stability were achieved by properly interacting the shock waves with the flameholding recirculation zone. The reason for the significant improvement in flame stability is believed to be the adverse pressure gradient caused by the shock, which can elongate the recirculation zone. Excessive shock strength (or poor shock placement) caused thermal choking to occur and the flame base moved upstream of the fuel tube exit. Optimization of the mixing and stability limits requires a careful matching of the shock-flame interaction location and the shock strength. The results show that the best mixing and stability corresponds to 10$\sp\circ$ wedges placed at an upstream position (4.0 d$\sb{F})$ such that the primary shocks create radial inflow near the flame base and interact with the recirculation zone. This upstream wedge position also allowed the second (recompression) set of shocks to provide radial inflow near the flame tip.
Detailed measurements of the velocity field in the symmetry plane of two jets and two
jet flames in a crossflow are obtained using particle image velocimetry. The jets issue
into a wind tunnel at density-weighted jet-to-crossflow velocity ratios r = 10 and r = 21,
with corresponding Reynolds numbers 6000 and 12 800. Ensemble statistics of the
velocity field are presented, and some interesting features of the entrainment process
in transverse jets are discussed. Deviations from the simple behaviour predicted by
the similarity analysis presented in Part 1 are highlighted. Simultaneous planar laser-induced
fluorescence imaging of the OH radical is performed in selected regions of the
flames. Results suggests that flame/flow interaction is strong near the lifted flamebase,
but increasingly weaker further downstream.
We present a similarity analysis of strong turbulent jets directed perpendicularly into
a crossflow. The analysis neglects pressure terms in the governing equations, and
assumes complete similarity in each of two intermediate-asymptotic regions of the
flow: a jet region, where the jet is largely unaffected by the crossflow, and a wake-like
region, where the jet has been deflected well into the crossflow. Scaling laws are derived
for velocity, scalar concentration, and jet trajectory, and show good agreement with
existing experimental data. The structure of the counter-rotating vortex pair implied
by this analysis is significantly different from typical representations found in the
literature.
This study assesses the prospect of main-fuel ignition with plasma-generating devices in a supersonic flow. Progress from this study has established baseline conditions for operation, such as the required operational time of a device to initiate a combustion shock train as predicted by computational fluid dynamics computations. Two plasma torches were investigated: a direct current constricted-arc design and an alternating current unconstricted-arc design based on a modified spark plug. Both plasma torches are realistic in size and operate within the same current and voltage constraints, although differing substantially in orifice geometry. to compare the potential of each concept, the flow physics of each part of the igniter/fuel-injector/combustor system was studied. To understand the constraints involved with the ignition process of a hydrocarbon fuel jet, an experimental effort to study gaseous and liquid hydrocarbons was conducted, involving the testing of ethylene and JP-7 fuels with nitrogen and air plasmas. Results from individual igniter studies have shown plasma igniters to produce hot pockets of highly excited gas with peak temperatures up to 5000 K at only 2 kW total input power. In addition, ethylene and JP-7 flames with a significant level of the hydroxyl radical, as determined by planar laser-induced fluorescence, were also produced in a Mach 2 supersonic flow with a total temperature and pressure of 590 K and 5.4 atm. Information from these experiments is being applied to the generation of constraints and the development of a configuration with perceived high ignition potential in full scramjet combustor testing.
A review of scramjet combustion simulation is provided in this paper. The topics covered include the fundamental problem of supersonic mixing layers, high-speed combustion modeling efforts, and actual calculations of realistic scramjet combustors. The review shows that the RANS approach dominates the turbulence modeling of the system, with only a handful of LES work. Also, virtually all the numerical works are based on low-order schemes, and the combustion models that have been used for realistic simulations solve the species evolution equations with assumed PDF closures, although there seems to be a growing use of the flamelet methods. Spray modeling for scramjet combustion has not received enough attention.
The difficulties with fueling of supersonic combustion ramjet engines with hydrocarbon based fuels presents many challenges. The need for a better solution to supersonic mixing has led to the development of many different styles of fuel injection. An aerodynamic ramp injector has been shown to have a quantitative improvement over a physical ramp while still achieving desirable mixing characteristics. Little quantitative combustion data is available on the performance of aerodynamic ramp injectors in a cavity-coupled scramjet flowpath, especially relative to round injectors. The objective of this study was to provide a direct comparison between the single angled injector and the aeroramp injector arrays in such a combustion environment. Ignition limits and pre-combustion shock position were used to define the operability differences while combustion efficiency and stream thrust were used for performance comparisons. These parameters were determined by operating a scramjet combustor in dual-mode operation over the range of Mach numbers expected during scramjet takeover from a boost vehicle. Performance and operability conclusions are based on raw data from static pressures, temperature measurements, and thrust stand loading. It was determined that the operability reduces significantly for the aeroramp injector for all conditions tested, but the performance is virtually identical to that of the round injectors. The aeroramp injector performance indicated improved near-field combustion with the potential for better performance in higher Mach number flow to include full supersonic combustion mode for which it was designed.