超声速湍流燃烧G/Z方程模型验证

为了研究超声速条件下的部分预混燃烧,引入一种基于Level set重构方法和稳态火焰面数据库的 G／Z方程模型,并利用德国宇航中心的DLR支板算例对 G／Z方程模型进行了验证。结果显示,超声速湍流燃烧G／Z方程模型可以捕捉到部分预混燃烧现象,数值模拟结果与实验结果吻合较好,验证了 G／Z方程模型运用到超声速部分预混条件下湍流燃烧流场计算的可行性。同时,超声速湍流燃烧 G／Z方程模型依赖于运用到的火焰传播速度模型与火焰面模型,模型精确度的提高有待进一步探究。     

天然气柔和燃烧锅炉结构参数影响模拟研究

在天然气锅炉中引入柔和燃烧技术将大大降低NOx排放,高速未燃气卷吸高温烟气回流并与之快速掺混再燃烧是柔和燃烧的重要特征,因此,开展天然气锅炉关键结构参数优化设计以组织流场形成柔和燃烧所需的高温低氧反应气氛非常必要。基于天然气锅炉的工况特征,设计了热负荷15kW的模型燃烧室,采用数值模拟手段详细研究了燃烧室高度、喷嘴孔径、喷嘴相对位置及烟气出口尺寸对燃烧室流场、组分场及关键参数——烟气回流比的影响规律,并最终确定了燃烧室结构优选方案,对天然气锅炉柔和燃烧机设计提供理论基础数据。     

高温低氧燃烧条件下氮氧化物的生成特性

高温低氧燃烧原理是高温空气燃烧技术赖以发展的基础，使得高温燃烧条件下的氮氧化物的生成与排放受到大大抑制。为了掌握这种非常规燃烧现象及污染物生成的基本规律，采用扩散燃烧模型、热力NO生成模拟与湍流N－S方程，数值研究了燃烧空间中空气氧浓度对燃烧特性和氮氧化物排放浓度的影响，再现了高温与低氧两种条件相结合，形成的稳定的低氮氧化物排放的燃烧特性。计算结果与实验数据吻合，为发展高温空气燃烧技术提供了理论基础。     

液体煤油超声速燃烧的三维数值分析

利用计算流体力学软件对煤油在双模态超声速燃烧室内的超声速喷雾燃烧进行数值模拟.采用离散液滴模型、概率密度函数紊流扩散燃烧模型和紊流k-ω模型,在飞行马赫数为5、煤油与空气的当量比为0.551时的超声速燃烧进行了计算.计算得到的壁面静压分布与实验数据十分接近;总压力损失系数是0.70,小于实验测量值0.707;分析数值结果可知,支板喷油和凹槽火焰稳定器提高了混合和燃烧效率,燃烧室出口燃烧效率达0.62,接近实验得到的燃烧效率0.696.     

燃烧参数影响高温空气燃烧特性的数值分析

高温空气燃烧技术具有高效节能和低NOx排放等多重优越性，是一种新型燃烧技术。为了深入研究高温空气燃烧机理和低氮氧化物排放特性，将湍流N—S方程与扩散燃烧模型和热力型NO生成模型相结合，研究了低氧浓度条件下，燃烧参数，如燃气供应量，过量空气系数，进口空气预热温度以及进口空气氧含量对燃烧的影响，为发展高温空气燃烧技术提供了理论依据。     

高温空气燃烧器换向过程非稳态流场和温度场的仿真分析

采用FLUENT软件和燃烧模型，对烧嘴交错布置的高温空气燃烧器换向后的非稳态过程进行了数值研究，换向后炉内的流场、温度场变化的计算结果表明，在换向后的前3S内燃烧炉的流场和温度场变化很大，但是经过3s的变化后，燃烧逐步稳定，最后重新形成稳定的燃烧，直至下一个换向前保持稳定燃烧。     

扁平射流煤粉燃烧器研究

扁平射流燃烧器是近年来我们发明的一种新型燃烧装置,它依据高温空气动力学原理,合理地组织燃烧室内空气动力场,达到了稳定火焰和高效燃烧的目的。本文对扁平射流燃烧室的结构、流场情况、气-固两相流场分布,以及热态实验和应用情况均进行了描述。在理论上给出了较合理的三区燃烧模型。实践表明应用扁平射流燃烧器有较大的经济效益。     

低速高温风洞设计及性能测试

设计了用于高温部件实验以及气膜冷却实验的低速直流回热风洞。该风洞由动力段、扩散段、稳定段、收缩段、实验段以及加热和回热设备构成,校核并测试了风洞的流场。利用实验段搭载的热电偶在常温工况下测定了风洞出口中心处1 h内的温升。在热态实验中将主流温度提高并利用热电偶测定实验段中心温度,建立校正关系。利用毕托管测定了风洞实验段常温工作的动压分布,计算了平均流速、平均动压以及流速和动压不稳定度。流场校验结果显示,风洞的流速以及动压稳定,常温工作下主流温升不明显,高温工况下温控性能卓越,适用于模拟高温流场、测量高温部件以及气膜冷却实验。     

温度条件对柴油机燃烧过程影响的研究

本文在模拟直喷柴油机的定容燃烧装置上，利用激光阴影像和燃烧火焰像同时成像的高速摄影方法研究了温度条件对柴油机燃烧过程的影响，实验结果表明，高温条件下火焰对柴油喷雾包围严重。     

二甲基醚燃烧过程及其优化的数学模拟（Ⅰ）——燃烧模型

在定容燃烧弹实验和发动机台架实验基础上建立了二甲基醚DME燃烧过程的准维多区的现象学燃烧模型，模型针对DME燃烧过程的特点，建立了新的喷雾混合和着火滞燃期子模型，修正了燃烧放热率子模型，研究了DME燃烧过程中氮氧化物（NOx)生成机理。模型的计算结果和实验结果相当吻合，模型对变工况、变参数有较好的适应能力。NOx生成历程计算分析表明，DME燃烧过程中NOx主要在扩散燃烧阶段生成，燃烧温度低是NOx排放低的主要原因。     

Numerical simulation of the interaction between shock train and combustion in three-dimensional M12-02 scramjet model

The characteristics of combustion flow fields and performance for hypersonic M12-02 scramjet were numerically simulated and analyzed. The compressible two-equation k-w SST turbulence model was employed for the turbulence model and the 9-species, 27-reaction-step hydrogen-air reaction mechanism was used as the reaction kinetics model. The numerical method was verified and a good agreement was obtained between the results of the numerical simulations and the experimental data. The results showed that shock waves from the upper and lower walls respectively crossed with each other near the central axis, forming a ‘diamond’ shape in the high-temperature combustion region. Compared to the conventional scramjet engine, most of the fuel reaction was in pure supersonic combustion mode for this hypersonic scramjet engine. Changes in the distribution of fuel on the upper and lower walls could have an appreciable impact on the combustion flow field. Average fuel distribution between upper and lower walls is benefit for combustion enhancement while the heat transfer in the corner of the side wall is severe and should be avoided during operation. The flame investigation showed that it cannot automatically predict the flame surface temperature in advance only based on the equivalence ratio Φ according to diffusion combustion theory. Compared to Φ = 1.0 condition, the flame surface temperature for Φ = 0.8 condition is higher as the complicated interaction between shock waves and combustion, which makes the local air temperature and mixing extent in flame surface is more appropriate. However, in terms of the overall engine performance, the Φ = 1.0 condition has the better combustion efficiency along the whole flow path.     

Numerical study on supersonic combustion of hydrogen and its mixture with Ethylene and methane with strut injection

In this paper, supersonic combustion and flow field of hydrogen and its mixture with ethylene and methane from strut injections in a Mach 2 supersonic flow are studied numerically. The fuel mixture of hydrogen, methane and ethylene represents the major products of pyrolysis of hydrocarbon fuels with large molecules such as kerosene as it acts as coolant and flows through cooling channels and absorbs heat. Detached Eddy Simulation with a reduced kinetic mechanism and steady flamelet model are applied to simulate turbulent combustion. The calculated temperature profiles of hydrogen are compared to the experimental results of DLR supersonic combustor for validation of the present numerical method. The supersonic combustion flows associated with shock waves, turbulent vortices and flame structures are studied. With addition of methane and ethylene, the flame zone moves further downstream of the strut and the maximum flow temperature at chamber exit decreases by 200 K. With analysis of total temperature ratios, it is found that combustion efficiency for hydrogen combustion is 0.91 and it decreases to 0.78 for the fuel mixture. The calculation of ignition delay time and flame speed reveals that fuel mixture of hydrogen and hydrocarbons has considerably larger delay time and smaller flame speed, that contributes to the weakened flame zone and lower combustion efficiency.     

Effects of plate vibration on the mixing and combustion of transverse hydrogen injection for scramjet

Transverse injection is an effective mixing enhancement technique for the combustor of scramjets. Vibration of the plate structure in combustor will easily be induced due to aerodynamic load and harsh aerothermodynamic load simultaneously. Effects of the plate vibration on the mixing and the combustion of the transverse hydrogen injection have been investigated numerically in this study. Finite rate chemistry model is used as combustion model. The supersonic jet experimental model of the Stanford University is modified slightly and used as the analysis model. Effects of the frequency and the amplitude of the plate vibration on combustion performance and flow field structure have been investigated in detail. The results show that the plate vibration increases the mixing efficiency, the combustion efficiency and the total pressure loss coefficient. Besides, it can change the flame structure and the shock wave structure, as well as increase the shock wave intensity at downstream of the injection. The vibration frequency has relatively little effect on the combustion efficiency and the total pressure loss coefficient. When the vibration frequency is large, it presents some high frequency pulsations for the total pressure loss coefficient. However, the vibration amplitude has large effect on combustion efficiency and the total pressure loss coefficient. When the vibration amplitude is small, the combustion efficiency presents regular periodic change with time. When the vibration amplitude is large, it diverges with time, and the flow tends to be unstable. The large vibration amplitude changes the stability of the original flow. Consequently, the combustion with large amplitude fluctuation can critically damage the combustion stability.     

Numerical investigation of wavy wall strut fuel injector for hydrogen fueled scramjet combustor

Mixing and combustion of a fuel with supersonic airstream in a scramjet combustor is a complex phenomenon because of very less resident time of the air in the combustion chamber. Mixing of fuel and air at supersonic speed and the subsequent combustion are greatly affected by the disturbance of the flow field in the form of shock waves, vortices and recirculation regions. In this research paper, the same concept has been considered by introducing an innovative strut fuel injector for the development of more shock waves and streamline vortices. The basic or standard computational domain of the scramjet combustor is considered from the reference of DLR experimental scramjet. The basic scramjet model consists of the wedge-shaped strut fuel injector. In this research, the strut injector has been re-designed such a way that to generate more oblique shock waves. Numerical analysis of the scramjet internal flow field has been performed with basic and innovative strut by solving the Reynolds-averaged Navier-Stokes equations with the help of computational fluid dynamics tool defined as ANSYS-FLUENT 16.0. The internal flow field of scramjet combustor with basic and innovative strut fuel injectors has been visualized from the analysis of pressure, temperature and velocity along with the analysis of flow structure, shock waves, and streamlines vortices. From the analysis of numerical results, it is identified that multiple numbers of oblique shock waves are being generated from the leading curved edge of the newly introduced strut. Both the pressure and temperature of airstream at the entrance of the combustion chamber are higher in the case of the wavy wall strut and it reduces the ignition delay time as compared to the basic strut model.     

Investigation of influence of detailed chemical kinetics mechanisms for hydrogen on supersonic combustion using large eddy simulation

Five detailed hydrogen combustion chemical kinetics mechanisms coupled with a partially stirred reactor (PaSR) combustion model were applied with large eddy simulation (LES) to study the influence of detailed mechanisms on supersonic combustion in a model scramjet combustor. The LES predictions of five detailed mechanisms for velocity, temperature, and combustor wall pressure show reasonable agreement with experimental results. Examining the effects on the distributions of temperature and species in supersonic combustion reveals that the supersonic flame structure is affected by detailed mechanisms. The different detailed mechanisms have a strong influence on the combustion efficiency, volume of the subsonic region, and subsonic combustion heat release rate in the combustor. Moreover, the total heat release in the computational domain for the five detailed chemical kinetics mechanisms is quite different. The subsonic combustion is dominant in the combustor for all detailed mechanisms. An analysis of the important reactions for H₂O, HO2, and OH is performed, revealing the reasons for differences in temperature and species distributions among the different detailed mechanisms in the combustor.     

Combustion modes of hydrogen jet combustion in a cavity-based supersonic combustor

Optical diagnosis-based combustion experiments were conducted to investigate the characteristics of cavity assisted hydrogen jet combustion in a supersonic flow with a total pressure of 1.6 MPa, a total temperature of 1486 K, and a Mach number of 2.52, simulating flight Mach 6 conditions. A supersonic combustor with a constant cross-sectional area was employed with several cavity configurations, fueling schemes and equivalence ratios. It was found that stable combustion could not be obtained without a cavity, indicating that pure jet-wake stabilized combustion could not be achieved and the cavity acted as a flameholder. Three combustion modes were observed for the cavity assisted hydrogen jet combustion: cavity assisted jet-wake stabilized combustion, cavity shear-layer stabilized combustion, and combined cavity shear-layer/recirculation stabilized combustion. The cavity assisted jet-wake stabilized combustion was observed to be the most unstable mode, accompanied by intermittent blowoff under the present conditions, while the combined cavity shear-layer/recirculation stabilized combustion mode seemed to be the most robust one.     

The initiation and propagation of detonation in supersonic combustible flow with boundary layer

Adaptive simulations solving the Navier-Stokes equations have been conducted in order to get a better understanding on the detonation initiation and propagation in a stoichiometric H₂/O₂/Ar supersonic mixture with boundary layer. The detonation is initiated by a continuous hot jet. When reflecting on the wall, the jet induced bow shock interacts with the boundary layer and forms the shock boundary layer interaction phenomena, while in Euler result the bow shock forms Mach reflection. The investigation shows that the Navier-Stokes simulation result is structurally in better agreement with the experiment compared with that of the inviscid Euler simulation result. The bow shock interacts with the separation shock, forming the shock induced combustion behind the interaction zone. Then the combustion front couples with shock and forms Mach stem induced detonation. The Mach stem induced detonation continues to getting higher and propagating upstream, initiating the main flow. The initiated partial detonation exists with the separation shock induced combustion front, forming an “oblique shock induced combustion-partial detonation” structure in the main flow. The investigation on the influence of free stream Mach number further confirms that the boundary layer has an important influence on detonation initiation. The parametric studies also show that there exists a free stream Mach number range to initiate the partial detonation in supersonic combustible flow successfully.     

Numerical investigations on the performance of a multistrut hydrogen injector in a scramjet combustor

This study aims at investigating the effect of a multistrut-based hydrogen injector in a scramjet combustor underreacting case. The numerical analysis is carried out using two-dimensional Reynolds-averaged Navier–Stokes equations with the Shear Stress Transport k − ω turbulence model in contention to comprehend the flow physics during scramjet combustion. The three major parameters, such as the shock wave pattern, wall pressures, and static temperature across the combustor, are validated with the reported experimental results. The results comply with the range, indicating that the adopted simulation method for single strut injection can be extended for other investigations. It is noticed that with multistrut injectors, as hydrogen jet pressure increases in the supersonic flow field, the jet penetration rate in the lateral direction of the flow and the total pressure loss as compared with the baseline injection pressure conditions has increased. The supersonic flow characteristics are determined based on the flow properties, combustion efficiency, mixing efficiency, and total pressure loss. Compared with the single-strut output of a scramjet combustor, multistruts inclusion increased the combustion efficiency by almost 18%, the mixing efficiency attained the maximum with 16% fewer lengths. The total pressure loss in single-strut is 14% lower than that of multistrut.     

Highly resolved numerical simulation of combustion in supersonic hydrogen–air coflowing jets

The present study is focused on the analysis of non-premixed combustion in high-velocity (supersonic) flows. The computations make use of a large eddy simulation (LES) model, which has been recently introduced to address combustion in high Reynolds number turbulent flows featuring moderate Damköhler values. We expect that the corresponding closure is able to account for the specificities encountered in high Mach number turbulent reactive flows featuring chemical reaction time scales with the same order of magnitude as flow time scales. The model takes finite-rate chemistry and micro-mixing effects into account within the framework of the partially stirred reactor (PaSR) concept, it is hereafter denoted by U-PaSR (unsteady partially stirred reactor). (i) In a first step of the present investigation, the capabilities of the U-PaSR closure hence proposed are evaluated through a detailed comparison performed between numerical results and the data obtained from an experimental study devoted to non-premixed combustion in supersonic co-flowing jets of hydrogen and vitiated air. The simulated test case corresponds to a well-documented experimental database that includes Raman scattering and laser-induced pre-dissociative fluorescence measurements. The comparisons performed between computational results and experimental data establish that the physical processes are well-described by the performed simulation. (ii) In a second step of this study, the flame structure and associated stabilization zone are analysed in the light of numerical simulation results. The post-processing to the computational results indeed confirms the importance of self-ignition processes, as well as the relevance of diagnostic tools recently introduced by Boivin et al. [1,2]. Considering the stabilization zone, it also emphasizes the essential importance of the pressure dynamics associated with the discharge of compressible coflowing jets into the atmosphere – an importance that was not so clearly evidenced from previous numerical simulations conducted on the same experimental benchmark.