碳氢燃料超声速燃烧分区实验研究

高保真度空天发动机数值模拟通常基于快速化学反应火焰面假设,即超声速燃烧反应的特征尺度小于湍流Kolmgorov尺度,该模型方法对于氢气燃料仿真计算结果较好,但对于乙烯等碳氢燃料仍需进一步研究.受限于极端环境特种非接触测量技术,目前尚未见超声速燃烧火焰分区判别的实验研究,导致目前超声速燃烧火焰面模型适用性以及分区燃烧物理模型认识不清,进而也制约了数值发动机技术发展.本工作基于自主研发的MHz发动机内窥光纤传感器,针对单边扩张双模态冲压发动机超声速燃烧火焰分区开展实验研究,通过化学自发光信号的最小香农熵定义超声速燃烧的特征时间τsc,根据理论方法和来流工况估算了超声速燃烧的流动特征时间,结合分区燃烧理论分析了双模态超燃冲压发动机内碳氢燃料燃烧的分区情况.通过燃烧分区情况以及与泰勒尺度的比较结果,验证了碳氢燃料超燃冲压发动机典型飞行条件下燃烧室内超声速燃烧处于旋涡小火焰区域(Re?50 000; Da∈1.80～2.60, B区),多尺度湍流涡结构发挥重要作用,并随着相对于泰勒尺度的不同大小,分别对应了不同尺度的涡结构主导该过程.同时给出了当量比、通量比以及来流马赫数对燃烧特征时间的影响规律...     

克服“高超声障''的途径”

在高超声速飞行条件下, 流入冲压发动机燃烧室并降至低速的空气温度, 随着飞行马赫数增 加升得愈来愈高. 燃料与高温空气混合燃烧释放的化学能将部分转化为解离能. 这些解离能 在长度受限的尾喷管中难以充分复合形成推力, 使冲压发动机性能随飞行马赫数增大而急剧 下降. 导致冲压发动机不适应高超声速飞行器的推进要求. 将此定名为``高超声障'. 半个 世纪以来, 广泛采用``超声速燃烧'降低流入燃烧室的空气温度来克服这种障碍. 虽已取得 不少进展, 然而关键性难点仍需继续攻克. 为了多途径促进吸气推进高超声速飞行的实现, 提出克服``高超声障'的另一种思路：保持现有冲压发动机吸气与燃烧方式, 通过催化促进 燃气解离组分在尾喷管膨胀过程中的复合, 增大冲压发动机的推力, 达到满足高超声速飞行 器的推进要求.     

高马赫数燃烧强化的激波风洞试验研究

基于中国科学院力学研究所的JF-24激波风洞, 通过开展高马赫数超燃冲压发动机的直连试验, 研究了高马赫数燃烧的强化方法以及燃料类型对燃烧的影响. 试验段是采用凹腔结构的圆截面燃烧室, 喷孔布置在隔离段, 燃料分别是氢气和乙烯, 当量比均为0.7. 燃料喷注分别采用无支板和小支板两种构型, 后者部分喷孔位于小支板顶部. 两种构型均设置了流向近距双排喷孔, 可分别进行单环和双环喷注. 试验结果论证了飞行马赫数10.0条件下氢气和乙烯在超高速气流中的稳定燃烧性能. 并且, 相比于单环喷注, 双环喷注以及补充小支板可以强化燃烧. 推测其原因是双环射流和激波/分离结构的近距离交互作用很可能改善掺混, 而补充小支板顶部喷注还能利用更多空气组织掺混. 在同样采用双环耦合小支板顶部喷注的强化措施下, 氢气与乙烯燃烧效率接近, 但氢推力性能更优. 这是因为较高热值氢的释热更多. 此外, 试验还证明了在当前来流条件下, 释热受控于掺混, 且高温离解效应限制释热上限. 这是由于释热降低流速且提高静温, 使高温离解的吸热效应更加显著.      

三维各向异性叉树网格流场自适应算法研究

建立了三维叉树形网格的数据结构,并将结构网格的有限体积法引入到叉树网格中,建立了相应的NS方程求解方法。在此基础上完善了包括各向异性自适应判别、合并/分裂、网格优化等步骤的算法,并提出了对流场结构进行“保护”性加密的优化加密方式。基于自适应叉树网格对高超声速横向喷流流场进行了数值模拟,捕捉到细致的流场结构,并将壁面压力系数计算值与文献试验值比较,得到了很好的模拟效果,具有较高的流场分辨精度。     

高超声速激波湍流边界层干扰直接数值模拟研究

高超声速激波与湍流边界层干扰会导致飞行器表面出现局部热流峰值,严重影响飞行器气动性能和飞行安全. 针对高马赫数激波干扰问题,以往数值研究多采用雷诺平均方法,而在直接数值模拟方面的相关工作较为少见. 开展高超声速激波与湍流边界层干扰的直接数值模拟研究,有助于进一步提升对其复杂流动机理认识和理解,同时也将为现有湍流模型和亚格子应力模型的改进提供理论依据. 采用直接数值模拟方法对来流马赫数6.0,34°压缩拐角内激波与湍流边界层的干扰问题进行了研究. 基于雷诺应力各向异性张量,分析了高超声速湍流边界层在压缩拐角内的演化特性. 通过对湍动能输运方程的逐项分析,系统地研究了可压缩效应对湍动能及其输运的影响机制. 采用动态模态分解方法,探讨了干扰流场的非定常运动历程. 研究结果表明,随着湍流边界层往下游发展,近壁湍流的雷诺应力状态由两组元轴对称状态逐渐演化为两组元状态,外层区域则由轴对称膨胀趋近于各向同性. 干扰流场内存在强内在压缩性效应(声效应),其对湍动能输运的影响主要体现在压力--膨胀项,而对膨胀--耗散项影响较小. 高超声速下压缩拐角内的非定常运动仍存在以分离泡膨胀/收缩为特征的低频振荡特性,其物理机制与分离泡剪切层密切相关.      

超燃冲压发动机燃烧模态转换试验研究

在模拟飞行高度为25 km、来流马赫数为6的情况下,采用试验研究的方法对超燃冲压发动机燃烧模态转换进行了直连式试验。根据燃烧室壁面压力分布和一维模型分析表明,燃料喷射位置和当量比的动态改变,实现了燃烧室内燃烧模态的动态转换。不同燃料喷射位置切换顺序比较表明,燃烧室内燃烧状态的改变受燃料分布所决定,但是燃烧室自身具有一定的抗波动能力。     

高超声速激波湍流边界层干扰直接数值模拟研究

高超声速激波与湍流边界层干扰会导致飞行器表面出现局部热流峰值,严重影响飞行器气动性能和飞行安全.针对高马赫数激波干扰问题,以往数值研究多采用雷诺平均方法,而在直接数值模拟方面的相关工作较为少见.开展高超声速激波与湍流边界层干扰的直接数值模拟研究,有助于进一步提升对其复杂流动机理认识和理解,同时也将为现有湍流模型和亚格子应力模型的改进提供理论依据.采用直接数值模拟方法对来流马赫数6.0,34?压缩拐角内激波与湍流边界层的干扰问题进行了研究.基于雷诺应力各向异性张量,分析了高超声速湍流边界层在压缩拐角内的演化特性.通过对湍动能输运方程的逐项分析,系统地研究了可压缩效应对湍动能及其输运的影响机制.采用动态模态分解方法,探讨了干扰流场的非定常运动历程.研究结果表明,随着湍流边界层往下游发展,近壁湍流的雷诺应力状态由两组元轴对称状态逐渐演化为两组元状态,外层区域则由轴对称膨胀趋近于各向同性.干扰流场内存在强内在压缩性效应(声效应),其对湍动能输运的影响主要体现在压力-膨胀项,而对膨胀-耗散项影响较小.高超声速下压缩拐角内的非定常运动仍存在以分离泡膨胀/收缩为特征的低频振荡特性,其物理机制与分离泡剪切层密切相关.     

斜爆震燃烧与斜爆震发动机研究进展

发展更高性能的吸气式高超动力成为未来高超声速飞行器研制的重中之重。现有基于煤油燃料的超燃冲压发动机,主要以爆燃模式组织燃烧,在高来流马赫数(Ma≥8)条件下,将面临高来流总温带来的高温离解和化学非平衡效应所带来燃料的能量难以充分释放和利用的难题,相比之下,斜爆震组织燃烧更接近于等容燃烧,具有燃烧释热速率快、热循环效率高等优势,是一种可应用于高马赫数吸气式动力的理想燃烧模式。斜爆震发动机能够显著缩短燃烧室长度,减少释热面积,是高马赫数飞行器极具潜力的吸气式动力方案。其中,斜爆震发动机内流道各部件的匹配设计、燃料喷注-混合、斜爆震波的起爆与驻定等是斜爆震发动机研制的关键技术,是当前高超声速领域的研究热点。但由于其面临的高速、高总温总压的来流条件以及爆震波在流场中的强间断与高速传播特性等,现有试验与数值模拟研究手段难以开展精细的燃烧流动机制研究,进而限制了相关控制机理的揭示与高精度模型的建立,使得斜爆震发动机工程研制较为困难,当前研究仍存在许多值得探讨的地方,文章在综述的同时对下一步研究提出相关建议。     

高焓湍流边界层壁面摩阻产生机制分析

高超飞行器在中低空以极高马赫数飞行时,飞行器表面会遇到湍流与高温非平衡效应耦合作用的新问题.这种高焓湍流边界层壁面摩阻产生机制是新型高超声速飞行器所关注的基础科学问题,厘清此产生机制可以为减阻方法的设计提供指导,具有重要的工程实用价值.本文选取高超声速飞行时楔形体头部斜激波后的高焓流动状态,开展了考虑高温非平衡效应的湍流边界层直接数值模拟研究,并设置同等边界层参数下的低焓完全气体湍流边界层流动作为对比,采用RD (Renard&Deck)分解技术研究了高焓湍流边界层摩阻的主要产生机制,对摩阻产生的主要贡献项积分函数分布进行了详细分析,研究了高温非平衡效应对摩阻产生的影响规律;采用象限分析技术,研究了摩阻分解湍动能生成项的主导流动事件.计算结果表明,高温非平衡效应会使得壁面摩阻脉动条带的流向和展向尺寸均减小.分子黏性耗散项和湍动能生成项是高焓湍流边界层摩阻生成的主要流动过程.分子黏性耗散项主要作用在近壁区,高焓流动的分布与低焓流动存在差异.象限分析表明,上抛和下扫运动是影响摩阻分解中湍动能生成项的主导事件.     

参考焓值法在高速槽道湍流中的修正

利用高来流马赫数为3, 5, 6, 7, 10的槽道湍流直接数值模拟(direct numerical simulations, DNS)数据, 评估和修正经典的参考焓值法. 研究表明在高来流马赫数槽道湍流中, 经典参考焓值法预测的壁面热流与DNS结果相差很大, 需要作适当的修正.修正参考焓值法Ⅰ和Ⅱ的预测结果明显优于经典参考焓值法;并且修正参考焓值法Ⅱ更加适用于高马赫数流动, 其壁面热流与DNS结果的相对误差在10%以内. 同时, 修正参考焓值法Ⅱ的普适性在超声速燃烧室隔离段热环境试验中得到了验证.     

Computations of a Hydrogen-Fueled Scramjet Combustor on Locally Refined Meshes

Simulations of an experimental hydrogen-fueled scramjet combustor are conducted using a novel dynamic hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation (DHRL) modeling framework. The combustor has a Mach 2 core flow with a ramp fuel injector resulting in an equivalence ratio of 0.17. Three grid resolutions are obtained using local refinement by a factor of two in each direction in the fuel mixing and combustion region, and results from the three grids are used to understand the effect of grid refinement. Simulations reproduce temperature, pressure, velocity, and fuel concentrations in reasonable agreement with experimental measurements. Although heat release decreases on average, as the mesh is refined, peaks of heat release are intensified causing locally elevated temperatures. Spectral analysis of turbulence kinetic energy and heat release suggests stringent resolution requirements for reacting simulations capable of accurately resolving the effects of chemical reactions. Using the medium grid the DHRL model is compared to the improved delayed detached eddy simulation (IDDES) model and two Reynolds-averaged Navier-Stokes (RANS) models. Overall, the DHRL framework significantly outperforms other methods when compared to the experimental pressure rise. Additionally, spectral analysis suggests that the current framework is capable of accurately resolving turbulent structures at frequencies higher than IDDES. The study is the first documenting the use of DHRL for supersonic reacting flow and results suggest that it is a viable alternative to existing turbulence treatments for these types of flows.     

Isothermal flow in a gas turbine combustor — a benchmark experimental study

An experimental investigation of the three-dimensional flow field within a water model of a can-type gas turbine combustion chamber is presented. Flow visualisation demonstrated that internal flow patterns simulated closely those expected in real combustors. The combustor comprised a swirl driven primary zone, annulus fed primary and dilution jets and an exit contraction nozzle. LDA measurements of the three mean velocity components and corresponding turbulence intensities were obtained to map out the flow development throughout the combustor. Besides providing information to aid understanding of the complex flow events inside combustors, the data are believed to be of sufficient quantity and quality to act as a benchmark test case for the assessment of the predictive accuracy of computational models for gas-turbine combustors.     

Performance of supersonic model combustors with staged injections of supercritical aviation kerosene

Supersonic model combustors using two-stage injections of supercritical kerosene were experimentally investigated in both Mach 2.5 and 3.0 model combustors with stagnation temperatures of approximately 1,750 K. Supercritical kerosene of approximately 760 K was prepared and injected in the overall equivalence ratio range of 0.5-1.46. Two pairs of integrated injector/flameholder cavity modules in tandem were used to facilitate fuel-air mixing and stable combustion. For single-stage fuel injection at an upstream location, it was found that the boundary layer separation could propagate into the isolator with increasing fuel equivalence ratio due to excessive local heat release, which in turns changed the entry airflow conditions. Moving the fuel injection to a further downstream location could alleviate the problem, while it would result in a decrease in combustion efficiency due to shorter fuel residence time. With two-stage fuel injections the overall combustor performance was shown to be improved and kerosene injections at fuel rich conditions could be reached without the upstream propagation of the boundary layer separation into the isolator. Furthermore, effects of the entry Mach number and pilot hydrogen on combustion performance were also studied.     

Subgrid Model Influence in Large Eddy Simulations of Non-reacting Flow in a Gas Turbine Combustor

Fuel efficiency improvement and harmful emission reduction are the paramount driving forces for development of gas turbine combustors. Lean-burn combustors can accomplish these goals, but require specific flow topologies to overcome their sensitivity to combustion instabilities. Large Eddy Simulations (LES) can accurately capture these complex and intrinsically unsteady flow fields, but estimating the appropriate numerical resolution and subgrid model(s) still remain challenges. This paper discusses the prediction of non-reacting flow fields in the DLR gas turbine model combustor using LES. Several important features of modern gas turbine combustors are present in this model combustor: multiple air swirlers and recirculation zones for flame stabilisation. Good overall agreement is obtained between LES outcomes and experimental results, both in terms of time-averaged and temporal RMS values. Findings of this study include a strong dependence of the opening angle of the swirling jet inside the combustion chamber on the subgrid viscosity, which acts mainly through the air mass flow split between the two swirlers in the DLR model combustor. This paper illustrates the ability of LES to obtain accurate flow field predictions in complex gas turbine combustors making use of open-source software and computational resources available to industry.     

OH-PLIF visualisation of radical farming supersonic combustion flows

Experimental investigations employing Planar Laser-induced fluorescence visualisation of the qualitative distribution of the OH radical (OH-PLIF), coupled with surface pressure measurements, have been made of flow in a generic, nominally two-dimensional inlet-injection radical farming supersonic combustion scramjet engine model. The test flows were provided by a hypersonic shock tunnel, and covered total enthalpies corresponding to the flight Mach number range 8.7–11.8 and approximately 150 kPa dynamic pressure. The surface pressure measurements displayed radical farming behaviour, that is a series of adjacent high and low pressure regions corresponding to successive shock/expansion structures, with no significant combustion-induced pressure rise until the second structure. OH-PLIF imaging between the first two structures provides the first direct experimental evidence of significant OH radical concentrations upstream of the ignition point in this mode of scramjet operation and shows that combustion reactions were occurring in highly localised regions in a complex turbulent and poorly micromixed fuel/air mixing layer confined to the fuel injection side of the combustor.     

斜爆轰发动机的推力性能理论分析

爆轰燃烧具有释热快、循环热效率高的特点. 斜爆轰发动机利用斜爆轰波进行燃烧组织, 在高超声速吸气式推进系统中具有重要地位. 以往研究主要关注斜爆轰波的起爆、驻定以及波系结构等, 缺少从整体层面出发对斜爆轰发动机开展推力性能分析. 本文将斜爆轰发动机内的流动和燃烧过程分解成进气压缩、燃料掺混、燃烧释热和排气膨胀4个基本模块并分别进行理论求解, 建立了斜爆轰发动机推力性能的理论分析模型. 在斜爆轰波系研究成果的基础上, 选取了过驱动斜爆轰、Chapman?Jouguet斜爆轰、过驱动正爆轰和斜激波诱导等容燃烧等4种燃烧模式来描述燃烧室内的燃烧释热过程, 并对比分析了不同燃烧模式对发动机比冲性能的影响. 此外, 还获得了不同来流参数、燃烧室参数和进排气参数等对发动机推力的影响规律, 发现来流马赫数和尾喷管的膨胀面积比是发动机理论燃料比冲的主要影响因素. 最后, 结合以往关于受限空间内斜爆轰波驻定特性等方面的研究成果, 提出了斜爆轰发动机燃烧室的设计方向.      

关于吸气式高超声速推进技术研究的思考

回顾了吸气式高超声速推进技术的研究进展, 分析了超燃冲压发动机研制面临的关键科学问题, 并从不同角度探讨了增大超燃冲压发动机推力的可能方法.这些方法包括: 能够降低总压损失的高超声速来流压缩方法、生成三维涡流的超声速混合增强技术、碳氢燃料的预热喷射、可以控制燃烧过程的燃烧室设计优化方法、通过减小发动机流道湿面积来降低摩擦阻力和催化复合解离的燃气降低高温气体效应.考虑到等压热力学循环的热效率,还建议研究在高超声速推进系统中应用热效率高的爆轰过程, 并探讨了爆轰推进方法研究的进展与问题.吸气式高超声速推进技术是高超声速飞行器发展的关键技术, 认真思考和探索其发展方向是非常必要的.      

AN ANALYTICAL DESIGN METHODOLOGY FOR AN ANNULAR COMBUSTOR

SUMMARY

This paper describes a computational procedure for the optimization of the performance parameters of a simulated annular combustor. This method has been applied to analyze the influence of the performance parameters and geometries on the annular combustor characteristics and provide a good understanding of combustor internal flow fields, and therefore it can be used for guiding the combustor design process. The approach is based on the solution of governing nonlinear, elliptic partial differential equations for 3-D axisymmetric recirculating turbulent reacting swirling flows and the modelling of turbulence, combustion, thermal radiation and pollutant formation. The turbulence effects are introduced via the modified two-equation κ-ε model. Turbulent combustion is modelled using the κ-ε-g model and a two-step turbulent combustion model is employed for the excess emission of carbon monoxide CO. For the evaluation of the NO pollutant formation rate, the NO pollutant formation model, which takes into account the influence of turbulence, presented here. The radiative heat transfer is handled by the heat flux model. The predictions of the combustor character-istics and performance parameters are made using the present approach.

Predictions of velocity, length of the recirculation zone, combustion efficiency and wall temperature are compared with measurements. Agreement between the predictions and experimental data is very satisfactory.     

微小尺度下氢气预混燃烧的实验研究

对内径为1.66mm的不锈钢管燃烧室的氢气预混燃烧实验进行了描述，采用 红外测温仪测量了燃烧室壁面的温度场分布，获得了不同燃烧热功率下的运行界限．在突扩 段内高温回流区的作用下，在带有5mm长突扩段的燃烧室内可以实现完全预混燃 烧，最高运行界限可达1.415．由于较高的进气速度和较大的燃烧室壁面散热，在不带突扩 段的不锈钢管内无法实现完全预混燃烧．结果表明突扩段对微小尺度燃烧具有稳定火焰、拓 宽燃烧运行界限的作用．通过对火焰形状和结构的观察，结合突扩段燃烧流场的分析，合理 解释了燃烧室壁面温度场随过量空气系数的变化规律．     

Rotating Detonation Combustor Research at the University of Cincinnati

Rotating detonation combustors (RDC) are at the forefront of pressure gain combustion (PGC) research. The simplicity in design and the ease of assembly makes it a promising technology that could be integrated into existing combustor architectures. This is, however, coupled with the considerable complexities of the detonation-based flow field, and the associated modes and coupling mechanisms. The current paper is an overview of the research done at the University of Cincinnati to address some of the challenges and questions pertaining to the physics of RDC operation. Issues such as combustor geometry, injection schemes and mixing, varied reactants behavior and modes of RDC operation are discussed. The effects of pressurization of the combustor, along with other detonation enhancement strategies are also deliberated upon. When appropriate, parallels are drawn to the phenomena of high frequency combustion instabilities to address the similarities in observations between the two fields.