纤维素、木质素对生物质与煤混烧特性的影响

利用热重红外分析仪着重分析了生物质中纤维素和木质素含量对混合热解及混烧特性的影响.生物质与煤的混合热解相比于纯煤的单一热解失重明显较大较快,并且纤维素含量越高,混合燃料失重越大;木质素含量较高,则混合燃料失重较慢.当燃烧温度为300～400℃之间时,混合燃料中纤维素含量越高,质量下降幅度越大;当燃烧温度为500～700℃之间时,混合燃料中则是木质素含量越高,质量下降幅度越大.     

支板凹腔一体化超燃冲压发动机实验研究

本文针对以凹腔支板一体化燃烧室为基本结构的超燃冲压模型发动机在自由射流风洞中的性能,主要研究了燃料在不同位置喷入时,燃烧室几何结构/气动性能/燃料混合及燃烧特性的相互耦合,以及对发动机推力性能的影响.结果表明支板与凹腔的一体化在合理配置燃料分布情况下可以获得较好的发动机性能.     

超细煤粉分级燃烧中NOx还原规律的研究

煤粉再燃技术脱氮效率高而运行费用低,是最行之有效的低NOx燃烧技术之一.通过模拟计算与试验方法,对一维热态煤粉炉内超细煤粉分级燃烧NOx的还原规律进行了研究.研究结果表明再燃燃料越细,对NOx的还原作用越强,最佳再燃燃料粒度为20μm;在相同NOx还原率的情况下,随着再燃燃料粒度的减小,需要的再燃燃料比例减小,再燃区停留时间缩短.以超细煤粉作为再燃燃料不仅使燃烧效率提高,而且对NOx的还原效率也相应提高,达70%左右.     

引射式甲醇燃烧器大负荷变动自适应配风特性

针对最优结构的甲醇混合燃料引射式燃烧器,完成了大范围负荷变动下自适应配风特性实验研究,获得了燃烧器的实际配风特性及燃烧特性,并验证了数值研究结果.结果表明:最佳结构的甲醇混合燃料引射式燃烧器可实现大范围负荷变动下燃烧器的自适应配风,且燃烧稳定.当燃烧器负荷从25%~120%变化时,摩尔引射系数随着燃烧器的负荷增加而少量减小,引射系数变化率为6.3%,燃烧效率能保持在99.2%以上。当燃烧器负荷小于40%时,燃烧温度随着燃烧器的负荷增加而增加,当负荷大于40%时,燃烧火焰温度基本稳定在1650 K。     

烧蚀支板对超燃冲压发动机燃烧室性能的影响

为了提高超燃冲压发动机燃烧室的性能,本文提出了燃料喷注支板与烧蚀支板组合的燃烧室新方案,并研究了新方案对超燃冲压发动机燃烧室性能的影响。相比于单燃料喷注支板方式而言,加入烧蚀支板后,虽然燃烧室内的总压恢复系数有所下降,但燃烧室内燃料与空气的混合效率、燃烧效率均有显著提高,燃烧效率的提高弥补了燃烧室内总压损失所带来的机械能损失,使得燃料喷注支板和烧蚀支板组合方式下的燃烧室比冲高于单燃料喷注支板时的比冲。     

污泥和准东煤的化学链燃烧实验研究

化学链燃烧是一种具有CO_2内分离的新型燃烧技术。基于天然铁矿石载氧体在1 kWth的串行流化床上进行了污泥和准东煤化学链混合燃烧实验。探究燃料反应器温度对碳转化率和反应器口体积分数的影响。在批次流化床上进行了固体燃料气化和热解实验。实验结果表明反应器温度由800到930℃,反应器出口的CO_2体积分数上升,CO和CH4降低,碳转化率升高。在整个温度范围内,相比于污泥,混合燃料对应的CO_2体积分数,但碳转化率低。在930℃时,混合燃料的碳转化率可以达到90%左右。准东煤中钠含量较高,但在连续运行过程中无烧结和团聚等问题出现,这主要归结于燃料混合导致的钠含量的降低以及高熔点钠化合物的生成。     

某支板式超音速剪切流燃烧的数值研究

针对某支板火焰稳定结构数值研究了二维超音速流动和燃烧规律,提出不同燃料供给方案,比较了采用全氢气、全甲烷和不同比例的混合燃气等情况下的燃烧性能.结果表明:单一燃料时,氢气超燃性能很好,但会出现热量雍塞,而甲烷无法燃烧,两种混合燃料方案均在燃烧室内出现了稳定的火焰,但氧气消耗率不理想,基于上述结论给出了一些提高超燃性能的改进措施.     

DMM燃料柴油机可控预混合燃烧的研究

本文介绍了在压燃式发动机上进行的预混合燃烧研究。在柴油机的进气道入口处安装了一个电控燃料喷射系统,喷入具有低十六烷值、低沸点的甲缩醛(DMM)燃料,在压缩冲程中形成均匀的混合气,并在上止点附近喷入少量柴油来点燃混合气。本文研究了预混合燃料比、发动机负荷、进气中CO2浓度和喷孔直径对发动机燃烧和排放的影响。试验结果表明,进气道喷射DMM的预混合燃烧能同时大幅降低NOx和碳烟排放,为降低柴油机有害排放提供了一种新途径。     

燃料空气供给条件对热声不稳定燃烧影响的研究

对Solar低排放预混燃烧系统的燃烧稳定性进行了数值研究.应用非定常N-S方程、雷诺应力紊流模型及涡团耗散燃烧模型,数值模拟了该类型燃烧器在不同的燃料空气供给条件下的气流流动特性和压力振荡特性,并给出了不稳定发生时压力和速度振荡的幅值和频率.根据供给条件的不同,燃烧可以是稳定的或是不稳定的,取决于燃料到火焰前沿的迟滞时间.采用CFD方法,可精确地获得燃料到火焰前沿的迟滞时间,证实了所采用的模型能够精确预测不稳定燃烧的出现及振荡特性.通过调整燃料与空气的供给条件,可使振荡激励或阻尼.     

惯性约束聚变中内爆混合的模型构建

从热斑质量方程和能量守恒方程入手,重新计算考虑混合后聚变燃料的比内能和比热容等热力学参数,分析混合效应在轫致辐射损失等能量输运方面的作用,构建有杂质混合情况的热斑燃烧动力学模型.根据静态模型中的热斑燃烧的功率平衡条件,研究烧蚀层杂质混合比例与点火阈值和热斑自持燃烧的关系.理论分析和数值计算表明,混合效应导致热斑中的轫致辐射增强是点火失败的重要因素之一.通过调整不同掺杂材料、混合浓度及混合方式,得到壳层混合与热斑面密度、热斑离子温度的演化之间的关系.最后,基于模拟结果给出两种降低混合影响的方法.     

Investigation on the flameholding mechanisms in supersonic flows: backward-facing step and cavity flameholder

Abstract  

As effective devices to extend the fuel residence time in supersonic flow and prolong the duration time for hypersonic vehicles cruising in the near-space with power, the backward-facing step and the cavity are widely employed in hypersonic airbreathing propulsive systems as flameholders. The two-dimensional coupled implicit RANS equations, the standard k-ε turbulence model, and the finite-rate/eddy-dissipation reaction model have been used to generate the flow field structures in the scramjet combustors with the backward-facing step and the cavity flameholders. The flameholding mechanism in the combustor has been investigated by comparing the flow field in the corner region of the backward-facing step with that around the cavity flameholder. The obtained results show that the numerical simulation results are in good agreement with the experimental data, and the different grid scales make only a slight difference to the numerical results. The vortices formed in the corner region of the backward-facing step, in the cavity and upstream of the fuel injector make a large difference to the enhancement of the mixing between the fuel and the free airstream, and they can prolong the residence time of the mixture and improve the combustion efficiency in the supersonic flow. The size of the recirculation zone in the scramjet combustor partially depends on the distance between the injection and the leading edge of the cavity. Further, the shock waves in the scramjet combustor with the cavity flameholder are much stronger than those that occur in the scramjet combustor with the backward-facing step, and this causes a large increase in the static pressure along the walls of the combustor.     

Simultaneous rainbow schlieren deflectometry and OH* chemiluminescence imaging of a diesel spray flame in constant pressure flow rig

The physical and chemical phenomena that take place during fuel injection, entrainment and fuel-air mixing, cool-flame and ignition reaction, and combustion in diesel sprays still require extensive study. Global parameters such as liquid and vapor jet penetration lengths and spreading rates render useful yet still limited information. Understanding of the temporal evolution of the spray as it progresses through various steps is needed to develop advanced clean combustion modes and high-fidelity predictive models with sufficient accuracy. In this study, high-speed rainbow schlieren deflectometry (RSD) and OH* chemiluminescence are used to simultaneously image fuel-air mixing, cool-flame reactions, ignition, flame propagation and stabilization, and combustion in a transient diesel-like flame. A constant pressure flow rig (CPFR) is used to conduct multiple injections in quick succession to obtain a statistically relevant dataset. n-heptane was injected at nominal supply pressure of 1000 bar from a single-hole diesel injector into ambient at pressure of 30 bar and temperature of 800 K. About 500 injections were performed and analyzed to reveal structural features of non-reacting and reacting regions of the spray, quantify jet penetration and spreading rates, and study cool-flame behavior, ignition, flame propagation and stabilization at lift-off length, and combustion at upstream and downstream locations.     

进口条件对超燃冲压发动机氢气燃烧流场特性的影响分析

对不同进口条件下的超燃冲压发动机燃烧室内氢气喷流超声速燃烧流动特性进行了数值模拟与分析.宽范围超燃冲压发动机是吸气式高超声速飞行器推进系统设计中的热点问题之一,受实验设备硬件条件及实验技术限制,数值模拟技术仍然是超燃冲压发动机燃烧室内燃气燃烧特性及流场特性的主要研究手段.采用基于混合网格技术的多组元N-S方程有限体积方法求解器,在不同进口Mach数及压强条件下,对带楔板/凹腔结构的燃烧室模型氢气喷流燃烧流场进行了数值模拟,对比分析了氢气喷流穿透深度、喷口前后回流区结构、掺混效率及燃烧效率等流场结构与典型流场参数的变化特性及影响规律.研究成果可为宽范围超燃冲压发动机喷流燃烧流动特性分析提供参考.      

新建高焓激波风洞Ma=8飞行模拟条件的实现与超燃实验

针对高Mach数超燃冲压发动机实验能力空缺问题,基于航天十一院新建的FD-21高能脉冲风洞,进行了Ma=8超燃飞行条件的模拟能力设计与调试,获得了总焓2.9 MJ/kg、总压11.01 MPa实验条件,实现了Ma=8、高度31 km飞行条件的风洞模拟.在此基础上,研发了匹配的氢燃料供应及喷注时序控制系统,设计了超燃冲压发动机模型,开展了超燃冲压发动机模型自由射流应用性风洞实验,获得了氢气燃料与空气、氮气超声速气流耦合流动作用下的实验模型壁面压力数据.在当量比近似一致条件下,空气来流对应的燃烧室壁面压力明显高于氮气来流情况,表明氢气在1 ms有效实验时间内完成了与超声速空气来流的混合、点火与燃烧,获得燃烧释热特性,确认了在FD-21高能脉冲风洞开展高Mach数超燃实验是切实可行的,为后续研究奠定了良好的基础.      

PSP measurement of flow fields in a supersonic combustor with a backward-facing step

The mixing fields within a SCRAM-jet combustion chamber are visualized using pressuresensitive paint (PSP) as an oxygen sensor. The experiments are performed in a small supersonic wind tunnel at the National Aerospace Laboratory — Kakuda Research Center (NAL-KRC). The main stream Mach number is 2.4, and the dynamic pressure ratios between the injected gas and the main flow are 0.3, 0.7, 1.1 and 1.5. Three fuel injection nozzles are used; oxygen is injected from the central nozzle and air from the two nozzles at either side. The spread of the injected gas is measured to observe the effects of placing the nozzles in different positions. The results show that the jet has its own independent flow structure, and that little mixing of gases occurs between the flow structures created by each nozzle. When the injection dynamic pressure ratio is increased, the oxygen fraction rises in the recirculation zone and falls in the separation zone downstream of the injection.     

Towards the development of an efficient low-NOx ammonia combustor for a micro gas turbine

Recent studies have demonstrated stable generation of power from pure ammonia combustion in a micro gas turbine (MGT) with a high combustion efficiency, thus overcoming some of the challenges that discouraged such applications of ammonia in the past. However, achievement of low NOx emission from ammonia combustors remains an important challenge. In this study, combustion techniques and combustor design for efficient combustion and low NOx emission from an ammonia MGT swirl combustor are proposed. The effects of fuel injection angle, combustor inlet temperature, equivalence ratio, and ambient pressure on flame stabilization and emissions were investigated in a laboratory high pressure combustion chamber. An FTIR gas analyser was employed in analysing the exhaust gases. Numerical modeling using OpenFOAM was done to better understand the dependence of NO emissions on the equivalence ratio. The result show that inclined fuel injection as opposed to vertical injection along the combustor central axis resulted to improved flame stability, and lower NH₃ and NOx emissions. Numerical and experimental results showed that a control of the equivalence ratio upstream of the combustor is critical for low NOx emission in a rich-lean ammonia combustor. NO emission had a minimum value at an upstream equivalence ratio of 1.10 in the experiments. Furthermore, NO emission was found to decrease with ambient pressure, especially for premixed combustion. For the rich-lean combustion strategy employed in this study, lower NOx emission was recorded in premixed combustion than in non-premixed combustion indicating the importance of mixture uniformity for low NOx emission from ammonia combustion. A prototype liner developed to enhance the control and uniformity of the equivalence ratio upstream of the combustor further improved ammonia combustion. With the proposed liner design, NOx emission of 42?ppmv and ammonia combustion efficiency of 99.5% were achieved at 0.3?MPa for fuel input power of 31.44?kW.     

Combustion characteristics of a dual-mode scramjet combustor with cavity flameholder

Combustion characteristics of a laboratory dual-mode ramjet/scramjet combustor were studied experimentally. The combustor consists of a sonic fuel jet injected into a supersonic crossflow upstream of a wall cavity pilot flame. These fundamental components are contained in many dual-mode combustor designs. Experiments were performed with an isolator entrance Mach number of 2.2. Air stagnation temperatures were varied from 1040 to 1490 K, which correspond to flight Mach numbers of 4.3–5.4. Both pure hydrogen and a mixture of hydrogen and ethylene fuels were used. High speed imaging of the flame luminosity was performed along with measurements of the isolator and combustor wall pressures. For ramjet mode operation, two distinct combustion stabilization locations were found for fuel injection a sufficient distance upstream of the cavity. At low T₀, the combustion was anchored at the leading edge of the cavity by heat release in the cavity shear layer. At high T₀, the combustion was stabilized a short distance downstream of the fuel injection jet in the jet-wake. For an intermediate range of T₀, the reaction zone oscillated between the jet-wake and cavity stabilization locations. Wall pressure measurements showed that cavity stabilized combustion was the steadiest, followed by jet-wake stabilized, and the oscillatory case. For fuel injection close to the cavity, a hybrid stabilization mode was found in which the reaction zone locations for the two stabilization modes overlapped. For this hybrid stabilization, cavity fueling rate was an important factor in the steadiness of the flow field. Scramjet mode combustion was found to only exist in the cavity stabilized location for the conditions studied.     

Effect of the incident shock wave interacting with transversal jet flow on the mixing and combustion

The interaction between an incident shock wave and a transverse jet flow for mixing and combustion in a supersonic airstream was investigated experimentally and numerically. NO planar laser induced fluorescence (NO-PLIF) and particle imaging velocimetry (PIV) for non-reactive flows and three-dimensional reactive/non-reactive numerical simulations were conducted to examine the effect of the incident shock wave on the three-dimensional flow structure and mixing mechanism between the airstream and the injected gas downstream of the injection slot. Results of NO-PLIF measurement and numerical simulation showed that, in the case without the incident shock wave, injected gas is seldom seen in the recirculation zone just downstream of the injection slot, while the injected gas with higher concentration is almost uniformly distributed in the recirculation zone when the incident shock wave is introduced downstream of the injection slot. Moreover, it was shown by the numerical simulations that the profiles of the local equivalence ratio is in the combustible range due to the enhanced entrainment of the airstream when the incident shock wave is introduced downstream of the injection slot. A large-scale recirculation in the direction parallel to the wall is generated by the three-dimensional flow effects, which enhances the mixing and extends the residence time in the recirculation zone in the case with incident shock wave downstream of the injection slot, the recirculation flow being confirmed successfully by PIV measurements as well. The results of three-dimensional reactive numerical simulations were in good agreement with the experimental flame-holding characteristics at a lower total temperature, which showed that flame-holding can be attained only when the incident shock wave was introduced downstream of the injection slot, confirming that the formation of three-dimensional and large-scale recirculation flow downstream of the injection slot enlarges the recirculation zone and enhances the mixing to produce the conditions for robust flame-holding.     

CFD analysis of the HyShot II scramjet combustor

The development of novel air-breathing engines such as supersonic combustion ramjets (scramjets) depends on the understanding of supersonic mixing, self-ignition and combustion. These aerothermochemical processes occur together in a scramjet engine and are notoriously difficult to understand. In the present study, we aim at analyzing the HyShot II scramjet combustor mounted in the High Enthalpy Shock Tunnel Göttingen (HEG) by using Reynolds Averaged Navier Stokes (RANS) and Large Eddy Simulation (LES) models with detailed and reduced chemistry. To account for the complicated flow in the HEG facility a zonal approach is adopted in which RANS is used to simulate the flow in the HEG nozzle and test-section, providing the necessary inflow boundary conditions for more detailed RANS and LES of the reacting flow in the HyShot combustor. Comparison of predicted wall pressures and heat fluxes with experimental data show good agreement, and in particular does the LES agree well with the experimental data. The LES results are used to elucidate the flow, mixing, self-ignition and subsequent combustion processes in the combustor. The combustor flow can be separated into the mixing zone, in which turbulent mixing from the jet-in-cross flow injectors dominates, the self-ignition zone, in which self-ignition rapidly takes place, and the turbulent combustion zone, located towards the end of the combustor, in which most of the heat release and volumetric expansion takes place. Self-ignition occurs at some distance downstream of the injectors, resulting in a distinct pressure rise further downstream due to the volumetric expansion as observed in the experiments. The jet penetration is about 30% of the combustor height and the combustion efficiency is found to be around 83%.     

The effects of mixture preburning on detonation wave propagation

Pressure gain combustion in the form of continuous detonations can provide a significant increase in the efficiency of a variety of propulsion and energy conversion devices. In this regard, rotating detonation engines (RDEs) that utilize an azimuthally-moving detonation wave in annular systems are increasingly seen as a viable approach to realizing pressure gain combustion. However, practical RDEs that employ non-premixed fuel and oxidizer injection need to minimize losses through a number of mechanisms, including turbulence-induced shock-front variations, incomplete fuel-air mixing, and premature deflagration. In this study, a canonical stratified detonation configuration is used to understand the impact of preburning on detonation efficiency. It was found that heat release ahead of the detonation wave leads to weaker shock fronts, delayed combustion of partially-oxidized fuel-air mixture, and non-compact heat release. Furthermore, large variations in wave speeds were observed, which is consistent with wave behavior in full-scale RDEs. Peak pressures in the compression region or near triple points were considerably lower than the theoretically-predicted values for ideal detonations. Analysis of the detonation structure indicates that this deflagration process is parasitic in nature, reducing the detonation efficiency but also leading to heat release far behind the wave that cannot directly strengthen the shock wave. This parasitic combustion leads to commensal combustion (heat release far downstream of the wave), indicating that it is the root cause of combustion efficiency losses.