Influence of jet-to-crossflow pressure ratio on nonreacting and reacting processes in a scramjet combustor with backward-facing steps

The jet-to-crossflow pressure ratio has a large impact on the combustion mode transition in the scramjet engine, and this information needs to be explored comprehensively. The effect of the jet-to-crossflow pressure ratio on the mixing and combustion processes in a backward-facing step combustor has been investigated numerically, and two similar cases have been utilized to validate the numerical approaches employed. The obtained results show that the wall pressure distribution for the nonreacting flow field has been predicted well, and the peak pressures are all a bit underestimated. However, the predicted wall pressure distribution for the reacting flow field does not match well with the experimental data, and it is overestimated. When the hydrogen is injected only from the bottom wall of the combustor, the mixing efficiency decreases with the increase of the jet-to-crossflow pressure ratio irrespective of the nonreacting or reacting flow field. When the hydrogen is injected simultaneously from the top and bottom walls, the separation shock wave is pushed forward to the entrance of the combustor, and it varies from an oblique one to a normal one. This means that the jet-to-crossflow pressure ratio has a great impact on the combustion mode transition for the scramjet engine, and the stable ramjet/scramjet mode transition can be obtained by controlling the fuel injection scheme.     

Bifurcation of flame structure in a lean-premixed swirl-stabilized combustor: transition from stable to unstable flame

The present work addresses unsteady flame dynamics in a lean-premixed swirl-stabilized combustor, with attention focused on the transition of flame structure from a stable to an unstable state. It was found that the inlet temperature and equivalence ratio are the two most important variables determining the stability characteristics of the combustor. A slight increase in the inlet mixture temperature across the stability boundary leads to a sudden increase in acoustic flow oscillation. One major factor contributing to this phenomenon is that as the inlet mixture temperature increases, the flame, which is originally anchored in the center recirculation zone, penetrates into the corner recirculation zone and flashes back, due to the increased flame speed. As a consequence, the flame is stabilized by both the corner- and the center-recirculating flows and exhibits a compact enveloped configuration. The flame flaps dynamically and drives flow oscillations through its influence on unsteady heat release. This problem has not previously been studied mechanistically. The results improve our understanding of the mechanisms of initiation and sustenance of combustion instabilities in gas-turbine engines with lean-premixed combustion.     

贫预混预蒸发燃烧室振荡燃烧特性的实验研究

针对燃用航空煤油的贫预混预蒸发模型燃烧室的振荡燃烧特性开展了实验研究。实验表明:在相同的燃烧室入口空气燃料混合物流速下,随着当量比的增加,燃烧室振荡燃烧的振荡主频从132 Hz增加到144 Hz,但燃烧室的均方根脉动压力幅值却从1 464 Pa下降到342 Pa。在当量比不变情况下,入流空气燃料混合物流速较低时,容易引发振荡燃烧现象,而当入流空气燃料混合物流速较高时,则燃烧会变得稳定。分析了整个燃烧实验装置的前4阶轴向声学模态频率,发现实验中所激励出的振荡燃烧主频和第二阶轴向声学模态频率吻合的很好。     

Swirling distributed combustion for clean energy conversion in gas turbine applications

Distributed combustion provides significant performance improvement of gas turbine combustors. Key features of distributed combustion includes uniform thermal field in the entire combustion chamber, thus avoiding hot-spot regions that promote NO_x emissions (from thermal NO_x) and significantly improved pattern factor. Rapid mixing between the injected fuel and hot oxidizer has been carefully explored for spontaneous ignition of the mixture to achieve distributed combustion reactions. Distributed reactions can be achieved in premixed, partially premixed or non-premixed modes of combustor operation with sufficient entrainment of hot and active species present in the flame and their rapid turbulent mixing with the reactants. Distributed combustion with swirl is investigated here for our quest to explore the beneficial aspects of such flows on clean combustion in simulated gas turbine combustion conditions. The goal is to develop high intensity combustor with ultra low emissions of NO and CO, and much improved pattern factor. Experimental results are reported from a cylindrical geometry combustor with different modes of fuel injection and gas exit stream location in the combustor. In all the configurations, air was injected tangentially to impart swirl to the flow inside the combustor. Ultra-low NO_x emissions were found for both the premixed and non-premixed combustion modes for the geometries investigated here. Swirling flow configuration, wherein the product gas exits axially resulted in characteristics closest to premixed combustion mode. Change in fuel injection location resulted in changing the combustion characteristics from traditional diffusion mode to distributed combustion regime. Results showed very low levels of NO (∼3 PPM) and CO (∼70 PPM) emissions even at rather high equivalence ratio of 0.7 at a high heat release intensity of 36 MW/m³-atm with non-premixed mode of combustion. Results are also reported on lean stability limit and OH^* chemiluminescence under both premixed and non-premixed conditions for determining the extent of distribution combustion conditions.     

Effect of flow field for colorless distributed combustion (CDC) for gas turbine combustion

Colorless distributed combustion (CDC) investigated here is focused on gas turbine combustion applications due to its significant benefits for, much reduced NOx emissions and noise reduction, and significantly improved pattern factor. CDC is characterized by distributed reaction zone of combustion which leads to uniform thermal field and avoidance of hot spot regions to provide significant improvement in pattern factor, lower sound levels and reduced NOx emission. Mixing between the combustion air and product gases to form hot and diluted oxidant prior to its mixing with the fuel is critical so that one must determine the most suitable mixing conditions to minimize the ignition delay. Spontaneous ignition of the fuel occurs to provide distributed reaction combustion conditions. The above requirements can be met with different configuration of fuel and air injections with carefully characterized flow field distribution within the combustion zone. This study examines four different sample configurations to achieve colorless distributed combustion conditions that reveal no visible color of the flame. They include a baseline diffusion flame configuration and three other configurations that provide conditions close to distributed combustion conditions. For all four modes same fuel and air injection diameters are used to examine the effect of flow field configuration on combustion characteristics. The results are compared from the four different configurations on flow field and fuel/air mixing using numerical simulations and with experiments using global flame signatures, exhaust emissions, acoustic signatures, and thermal field. Both numerical simulations and experiments are performed at a constant heat load of 25 kW, using methane as the fuel at atmospheric pressure using normal temperature air and fuel. Lower NOx and CO emissions, better thermal field uniformity, and lower acoustic levels have been observed when the flame approached CDC mode as compared to the baseline case of a diffusion flame. The reaction zone is observed to be uniformly distributed over the entire combustor volume when the visible flame signatures approached CDC mode.     

Analysis of combustion instability of hydrogen fueled scramjet combustor on high-speed OH-PLIF measurements and dynamic mode decomposition

This paper investigated the combustion instability of spanwise positions in a hydrogen fueled scramjet combustor with a cavity flame holder. High-speed OH-PLIF technique was performed on a direct-connect supersonic combustion facility, and dynamic mode decomposition (DMD) as postprocessing. Combustion instability was investigated by characterizing the dominant frequencies and growth factors. By changing the equivalence ratio of hydrogen, the peak frequencies of scramjet mode and ramjet mode were obtained. Scramjet mode tended to have small oscillation at 150–200 Hz reflected by negative growth factors due to the stable flame structure. At ram-to-scram transition, oscillations at 80–120 Hz were remarkably enhanced due to the positive growth factors. In ramjet mode, the large differences of frequency characteristics in spanwise positions were observed. The dominant DMD modes near the cavity wall appeared to have negative growth factors leading to a stable flame with small oscillations. Besides, the characteristics of frequency-shift were affected by the positions of injector.     

Investigation on combustion mode and heat release in a model scramjet engine affected by shocks

A model scramjet engine in which the 1.0 Ma hydrogen jet mixes and reacts with the 2.0 Ma surrounding airstream is investigated using large eddy simulation. The flame structure is analyzed with a focus on the relationship between premixed/diffusion combustion mode and heat release in the supersonic reacting flow. The flame filter is used to evaluate the contributions to heat release rate by different combustion modes qualitatively and quantitatively. Results show that the heat is released from a combination of premixed combustion mode and diffusion combustion mode even when the fuel and airstream are injected into the combustor separately. Local mode-transition occurs as the supersonic jet flame propagates and interacts with shocks. The diffusion combustion mode dominates during the ignition stage and the premixed combustion becomes dominant during the intensive combustion region. When the shock wave impinges on the flame, the combustion area decreases a little due to the compression effects of the shock. However, the heat release rate is significantly improved in the interaction region since the shock could increase the air entrainment rate by directing the airflow toward the fuel jet and enhance the mixing rate by inducing vorticity due to baroclinic effects, which is good for flame stabilization in the supersonic flow. For the present case, 33.3% of the heat is released by diffusion combustion and 66.7% of the heat is released by premixed combustion. Thus the premixed combustion mode is dominant in terms of its contributions to heat release in the model scramjet engine.     

On the phase between pressure and heat release fluctuations for propane/hydrogen flames and its role in mode transitions

This paper presents an experimental investigation into mode-transitions observed in a 50-kW, atmospheric pressure, backward-facing step combustor burning lean premixed C₃H₈/H₂ fuel mixtures over a range of equivalence ratios, fuel compositions and preheat temperatures. The combustor exhibits distinct acoustic response and dynamic flame shape (collectively referred to as “dynamic modes”) depending on the operating conditions. We simultaneously measure the dynamic pressure and flame chemiluminescence to examine the phase between pressure (p′) and heat release fluctuations (q′) in the observed dynamic modes. Results show that the heat release is in phase with the pressure oscillations (θ_qp ≈ 0) at the onset of a dynamic mode, while as the operating conditions change within the mode, the phase grows until it reaches a critical value θ_qp = θ_c, at which the combustor switches to another dynamic mode. According to the classical Rayleigh criterion, this critical phase (θ_c) should be π/2, whereas our data show that the transition occurs well below this value. A linear acoustic energy balance shows that this critical phase marks the point where acoustic losses across the system boundaries equal the energy addition from the combustion process to the acoustic field. Based on the extended Rayleigh criterion in which the acoustic energy fluxes through the system boundaries as well as the typical Rayleigh source term (p′q′) are included, we derive an extended Rayleigh index defined as R_e = θ_qp/θ_c, which varies between 0 and 1. This index, plotted against a density-weighted strained consumption speed, indicates that the impact of the operating parameters on the dynamic mode selection of the combustor collapses onto a family of curves, which quantify the state of the combustor within a dynamic mode. At R_e = 0, the combustor enters a mode, and switches to another as R_e approaches 1. The results provide a metric for quantifying the instability margins of fuel-flexible combustors operating at a wide range of conditions.     

多流体模型在燃气轮机燃烧室三维反应流中的应用

对一种模型燃气轮机燃烧室中的三维反应流进行了数值模拟，模型燃烧室的燃料是CH4，燃烧类型是预混燃烧，在数值模拟过程中，采用了Spalding于1995年提出来的多流体模型来对燃烧室中的湍流预混燃烧进行了数值模拟，在数值模拟过程中考虑了辐射问题，采用了六通量辐射模型。通过数值模拟给出了速度，压力，湍流脉动动能，湍流动能耗散率，焓值，湍流粘度，温度，密度，燃烧产物质量分数，氧的质量分数，燃料/空气混合比，燃料质量分数，空间三个方向的辐射热通量以及各种流体的质量分数等变量的分布情况，此外，还采用传统的旋涡破碎模型对此燃烧室进行了数值模拟，并对两种方法的结果进行了分析比较，由分析可以看出多流体模型的结果接近于实际情况，对模型燃烧室进行三维反应流数值模拟的工作为今后对实际燃烧室反应流的数值模拟打下了一定的基础。     

Numerical investigations of mixed supersonic and subsonic combustion modes in a model combustor

Flame dynamics and statistics of mixed supersonic and subsonic combustion modes under different air inflow and global equivalence ratio conditions in a hydrogen-fueled model combustor are numerically studied. Three methods including spanwise-averaged Mach number, spanwise-averaged Mach number conditioning on the local heat release, and fraction of heat release are proposed to identify supersonic and subsonic combustion modes. The probability distributions of supersonic and subsonic combustion modes are also analyzed based on the statistics on multiple instantaneous snapshots of the numerical results. The critical global equivalence ratio for thermal choking in a range of supersonic inflow conditions is derived theoretically based on a one-dimensional duct flow with heat addition. Furthermore, it is found that the flame lift-off distance in both supersonic and subsonic flows decreases with increased air inflow velocity, but increases with global equivalence ratio. The fraction of supersonic heat release and its oscillation increase with increased air inflow velocity.     

Combustion modes periodical transition in a hydrogen-fueled scramjet combustor with rear-wall-expansion cavity flameholder

Cavity-stabilized combustion modes periodical transition in a laboratory hydrogen-fueled scramjet combustor were captured by high speed imaging. Experiments were performed with an isolator entrance Mach number of 2.92, and a sonic transverse fuel jet upstream of the cavity was employed. The reproducibility of the results had been tested in several repeated experiments. It is observed that, under a fixed combustor entry flow parameters and fuel equivalence ratio, after the flame has been fully developed, the flame structure may periodically transform between two different combustion modes. Specifically, under a moderate equivalence ratio, the flame structure switches between the cavity shear-layer stabilized mode and the combined shear-layer/recirculation stabilized mode, which shows an apparent periodicity with a period of about 7.5 ms. The formation mechanism for this novel phenomena is analyzed and some suppositions are given. When increasing the equivalence ratio to a high level, the flame structure shows a quasi-periodic low-frequency oscillation and the combustion mode changes between the combined shear-layer/recirculation stabilized mode and the cavity-assisted jet wake stabilized mode. At last, the formation mechanism and characteristics of the combined shear-layer/recirculation stabilized mode are detailed analyzed.     

Hydrogen-fueled detonation ramjet model: Wind tunnel tests at approach air stream Mach number 5.7 and stagnation temperature 1500 K

The mode of continuous spinning detonation (CSD) combustion of hydrogen in the annular combustor of a model of a hydrogen-fueled detonation ramjet under conditions of approach air stream Mach number 5.7 and stagnation temperature 1500 K is registered experimentally in a short-duration (pulsed) wind tunnel at the overall air-to-hydrogen equivalence ratio (ER) ranging from 0.7 to 1.4. The maximum values of thrust and specific impulse of the ramjet model are attained at ER = 1.25 and are estimated as 1550 N and 3300 s, respectively. At 1.4 < ER < 1.6, the mode of longitudinally pulsating detonation (LPD) combustion is registered with somewhat lower values of thrust and specific impulse.     

Investigation of flow characteristics in supersonic combustion ramjet combustor toward improvement of combustion efficiency

Supersonic combustion ramjet (scramjet) is a variant of ramjet in which the combustion takes place at supersonic velocity. The flow physics inside the scramjet combustor is quite complex due to the fact that the mixing and completion of the combustion take place in a short time, which is of the order of milliseconds. This study focuses on flow characteristics within the combustion chamber of the scramjet engine that is designed to improve energy efficiency by enhancing combustion efficiency. The effect on combustion performance and thereby the energy efficiency on using strut‐based flame stabilizer is evaluated at different positions. Reynolds averaged Navier‐Stokes equations are solved with the Shear Stress Transport k‐ω turbulence model. Single strut configuration is used to validate with the experimental data. Single strut is then compared with three‐strut configuration. In the three‐strut configuration, the location of the primary strut is kept constant, and the secondary struts are relocated in x and y directions. Combustion performance was evaluated for the cases of flow from primary strut only and through three struts. It was found that the placement of secondary strut in a three‐strut configuration plays a vital role in improving energy efficiency. A maximum of 33.86% improvement in combustion efficiency was observed in comparison to the single strut combustor. A reduction in unburned fuel was observed, making the system more energy efficient. If the struts are not placed optimally, the combustion performance of the combustor was observed to be lower than that of a single‐strut configuration. The shock reflection and expansion fans within the primary combustion zone and the secondary strut region enhance the combustion efficiency. The wall static pressure was observed to increase with the addition of secondary struts. For certain strut configurations, flow separation was seen on the combustor walls. If the secondary strut was placed close to the primary strut, combustion efficiency was found to enhance. It was seen that combustion efficiency was also enhanced for the cases of reacting flow from primary strut only. It could also help to increase fuel efficiency, as additional fuel is not supplied to the secondary strut, making the overall system energy efficient. As the secondary strut is introduced, total pressure loss also increases. It could also be noted that if the combustor length was increased, there could be further increased in combustion efficiency.     

Transverse combustion instabilities: Acoustic,fluid mechanic,and flame processes

Thermoacoustic oscillations associated with transverse acoustic modes are routinely encountered in combustion chambers. While a large literature on this topic exists for rockets, no systematic reviews of transverse oscillations are available for air-breathing systems, such as in boilers, aircraft engines, jet engine augmentors, or power generating gas turbines. This paper reviews work on the problem for air-breathing systems, summarizing experimental, modeling, and active control studies of transverse oscillations. It then details the key physical processes controlling these oscillations by describing transverse acoustic wave motions, the effect of transverse acoustic waves on hydrodynamic instabilities, and the influence of acoustic and hydrodynamic fluid motions on the unsteady heat release. This paper particularly emphasizes the distinctions between the direct and indirect effect of transverse wave motions, by arguing that the dominant effect of the transverse acoustics is to act as the “clock” that controls the frequency and modal structure of the disturbance field. However, in many instances, it is the indirect axial flow disturbances at the nozzles (driven by pressure oscillations from the transverse mode), and the vortices that they excite, that cause the dominant heat release rate oscillations. Throughout the review, we discuss issues associated with simulating or scaling instabilities, either in subscale experimental geometries or by attempting to understand instability physics using identical nozzle hardware during axial oscillations of the same frequency as the transverse mode of interest. This review closes with a model problem that integrates many of these controlling elements, as well as recommendations for future research needs.     

Investigation of reverse flow distributed combustion for gas turbine application

In this paper reverse flow modes of colorless distributed combustion (CDC) have been investigated for application to gas turbine combustors. Rapid mixing between the injected fuel and hot oxidizer has been carefully explored for spontaneous ignition of the mixture to achieve distributed combustion reactions. Distributed reactions can be achieved in premixed, partially premixed or non-premixed modes of combustor operation with sufficient entrainment of burned gases and faster turbulent mixing between the reactants. In the present investigation reverse flow modes consisting of three configurations at thermal intensity of 28 MW/m³-atm and five configurations at thermal intensity of 57 MW/m³-atm have been investigated and these high thermal loadings represent characteristic gas turbine combustion conditions. In all the configurations the air injection port is positioned at the combustor exit end, whereas the location of fuel injection ports is changed to give different configurations. The results are presented on the exhaust emissions and radical emissions using experiments, and evaluation of flowfield using numerical simulations. Ultra-low NO_x emissions were found for both the premixed and non-premixed combustion modes investigated here. Cross-flow configuration, wherein the fuel is injected at high velocity cross stream to the air jet resulted in characteristics closest to premixed combustion mode. Change in fuel injection location resulted in changing the combustion characteristics from closer to diffusion mode to distributed regime. This feature is beneficial for part load operation where higher stability limit is desirable.     

Analysis of unsteady reacting flows and impact of chemistry description in Large Eddy Simulations of side-dump ramjet combustors

Among all the undesired phenomena observed in ramjet combustors, combustion instabilities are of foremost importance and predicting them using Large Eddy Simulation (LES) is an active research field. While acoustics are naturally captured by compressible LES provided that the proper boundary conditions are applied, combustion/chemistry modelling remains a critical issue and its impact on numerical predictions must still be assessed for complex applications. To do so, two different ramjet LES’s are compared here. The first simulation is based on a standard one-step chemistry known to over-estimate the laminar flame speed in fuel rich conditions. The second simulation uses the same scheme but introduces a correction of reaction rates for rich flames to match a detailed mechanism provided by Peters (1993) [1]. Even though the two chemical schemes are very similar and very few points burn in rich regimes, distinct limit-cycles are obtained with LES depending on which scheme is used. Results obtained with the standard one-step chemistry exhibit high frequency self-sustained oscillations. Multiple flame fronts are stabilized in the vicinity of the shear layer developing at the exit of the air inlets. When compared to the experiment, the fitted one-step scheme yields better predictions than the standard scheme. With the fitted scheme, the flame is detached from the air inlets and stabilizes in the regions identified in the experiment (Ristori et al. (2005) [2], Heid and Ristori (2003) [3], Heid and Ristori (2005) [4], Ristori et al. (1999) [5]). LES and experiments exhibit all main low-frequency modes including the first longitudinal acoustic mode. The high frequencies excited with the standard scheme are damped with the fitted scheme. The chemical scheme is found, for this ramjet burner, to have a strong impact on the predicted stability: approximate chemical schemes even in a limited range of equivalence ratio can lead to the occurence of non-physical combustion oscillations.     

An experimental study of fuel–air mixing section on unstable combustion in a dump combustor

The main objective of this study is effect of the various fuel–air mixing section geometries on the unstable combustion. For the purpose of observing the combustion pressure oscillation and phase difference at each of the dynamic pressure results, the multi-channel dynamic pressure transducers were located on the combustor and inlet mixing section. By using an optically accessible quartz-type combustor, we were able to OH* measurements to characterize the flame structure and heat release oscillation with the use of a high-speed ICCD camera. In this study, we observed two dominant instability frequencies. Lower frequencies were measured around 240–380 Hz, which were associated with a fundamental longitudinal mode of combustor length. Higher frequencies were measured around 410–830 Hz. These were related to the secondary longitudinal mode in the combustion chamber and the secondary quarter-wave mode in the inlet mixing section. These second instability mode characteristics are coupled with the conditions of the combustor and inlet mixing section acoustic geometry. Also, these higher combustion instability characteristics include dynamic pressure oscillation of the inlet mixing section part, which was larger than the combustor section. As a result, combustion instability was strongly affected by the acoustically coupling of the combustor and inlet mixing section geometry.     

Impact of parametric variation on combustion characteristics of hydrogen-fueled strut based scramjet combustor at supersonic speed

Supersonic combustion ramjet engine is more fascinating among all the air-breathing engines. Due to its higher thrust to weight ratio, researchers are more interested to get the superior combustion performance at the optimum boundary conditions. The flow field characteristics and combustion performance have been analysed with the help of Ansys 14.0 software. Generic scramjet combustor of German Aerospace Center (DLR) has been taken into consideration for comparison purpose and off design analysis has been conducted to investigate and analyse the changes. Two dimensional compressible Reynolds Averaged Navier-Stokes (RANS) turbulence model has been opted with the finite-rate/eddy-dissipation reaction model. K-ε two equation turbulence model has been selected to reach up to reasonable accuracy. Validation of the present work has been done with the help of both non-reacting and reacting type data from open literature. To choose the appropriate meshing of the computational model three different types of mesh elements, that is, coarse, medium and fine has been analysed and also grid independence analysis is performed. The present article objective is to get optimum boundary condition by changing the incoming air temperature and pressure at constant Mach number to connect the bridge between incoming air temperature and pressure to the change in velocity throughout the combustion chamber. The detailed understanding and explanation have been done by varying the temperature range of incoming air because of its major impact on combustion performance. Nonetheless, a small variation of air pressure will also discuss to observe the parameters which majorly influence while doing performance analysis. At the end the Optimum boundary condition for the present computation work is observed to be at 833 K temperature with 115 299 Pa pressure.     

Combustion dynamics of a low-swirl combustor

The methodology for the measurement of dynamic combustor behavior has never been clearly established, due to the complexities associated with unsteady premixed flames and the difficulties in their measurement. The global and local distribution of Rayleigh index and the flame response functions are the main parameters normally employed to quantify and describe combustion dynamics. The Rayleigh index quantifies the thermoacoustic coupling, while the flame response function is a measure of the response of the system to outside disturbances. The primary objective of this work is to investigate the combustion dynamics of a commonly used low-swirl burner and to develop tools and methods for examining the dynamics of a combustion system. To this end, the effect of acoustic forcing at various frequencies on flame heat release behavior has been investigated. The current work uses OH-PLIF imaging of the flame region to produce phase-resolved measurements of flame behavior at each frequency. The response of the flame to the imposed acoustic field over the range of 22-400 Hz is then calculated from the processed images. This provides a starting point for an extension/extrapolation to practical acoustic ranges (∼5000 Hz). It was found that the thermoacoustic coupling was mainly evident in the shear mixing zone, producing a toroidal Rayleigh index distribution pattern. The phase shift of the flame fluctuation from the imposed acoustic wave seems to be very closely coupled to the vortices generated at the flame boundary due to shear mixing (Kelvin-Helmholtz instability), thus inducing the alternating toroidal structures. The peak value of the flame response function coincides with the peak absolute value of the Rayleigh index.