关于超声速燃烧与高超动力

先进发动机是航空工业的核心技术,而吸气式高超声速发动机一直是宇航飞行技术研发的首位难题.发动机的性能依赖于其能量转换模式和燃烧组织方法,相关理论研究具有基础性和启发性意义.论文首先讨论了超声速燃烧,它一直是超燃冲压发动机技术的理论基础.然后综述了相关研究进展,提出了吸气式高超声速冲压推进技术的3个临界条件,或者称为临界...     

Supersonic combustion and hypersonic propulsion

50 多年的努力和曲折经历证明了超声速燃烧冲压发动机概念的可行性. 本文对影响超燃冲压发动机技术成熟的主要因素作了扼要的分析. 高超声速推进的首要问题是净推力, 利用超声速燃烧获得推力遇到各种实际问题的制约, 它们往往互相牵制. 几次飞行试验表明高超声速飞行需要的发动机净推力仍差强人意, 液体碳氢燃料(煤油) 超燃冲压发动机在飞行马赫数5 上下的加速和模态转换过程, 成为高超声速吸气式推进继续发展的瓶颈. 研究表明, 利用吸热碳氢燃料不仅是发动机冷却的需要也是提高发动机推力和性能的关键举措, 燃料吸热后物性改变对燃烧性能的附加贡献对超燃冲压发动机的净推力至关重要.当前, 实验模拟技术和测量技术相对地落后, 无法对环境、尺寸和试验时间做到完全的模拟. 计算流体动力学(Computational Fluid Dynamics, CFD) 逐渐成为除实验以外唯一可用的工具, 然而, 超声速燃烧的数值模拟遇到湍流和化学反应动力学的双重困难. 影响对发动机的性能作正确可靠的评估.提出双模态超燃冲压发动机模态转换、吸热碳氢燃料主动冷却燃料催化裂解与超声速燃烧耦合、燃烧稳定性、实验模拟技术与装置、内流场特性和发动机性能测量、数值模拟中的湍流模型、煤油替代燃料及简化机理等研究前沿课题, 和未来5~10 年重点发展方向的建议.     

Functional mathematical model of a hydrogen-driven combustion chamber for a scramjet

A functional mathematical model of a hydrogen-driven combustion chamber for a scramjet is described. The model is constructed with the use of one-dimensional steady gas-dynamic equations and parametrization of the channel configuration and the governing parameters (fuel injection into the flow, fuel burnout along the channel, dissipation of kinetic energy, removal of some part of energy generated by gases for modeling cooling of channel walls by the fuel) with allowance for real thermophysical properties of the gases. Through parametric calculations, it is found that fuel injection in three cross sections of the channel consisting of segments with weak and strong expansion ensures a supersonic velocity of combustion products in the range of free-stream Mach numbers M_∞ = 6–12. It is demonstrated that the angle between the velocity vectors of the gaseous hydrogen flow and the main gas flow can be fairly large in the case of distributed injection of the fuel. This allows effective control of the mixing process. It is proposed to use the exergy of combustion products as a criterion of the efficiency of heat supply in the combustion chamber. Based on the calculated values of exergy, the critical free-stream Mach number that still allows scramjet operation is estimated.     

Large Eddy simulation of turbulent hydrogen-fuelled supersonic combustion in an air cross-flow

The main aim of this article is to provide a theoretical understanding of the physics of supersonic mixing and combustion. Research in advanced air-breathing propulsion systems able to push vehicles well beyond $M=4$ M = 4 is of interest around the world. In a scramjet, the air stream flow captured by the inlet is decelerated but still maintains supersonic conditions. As the residence time is very short $(\sim \!\!\mathrm{1ms})$ ( ～ 1 ms ) , the study of an efficient mixing and combustion is a key issue in the ongoing research on compressible flows. Due to experimental difficulties in measuring complex high-speed unsteady flowfields, the most convenient way to understand unsteady features of supersonic mixing and combustion is to use computational fluid dynamics. This work investigates supersonic combustion physics in the Hyshot II combustion chamber within the Large Eddy simulation framework. The resolution of this turbulent compressible reacting flow requires: (1) highly accurate non-dissipative numerical schemes to properly simulate strong gradients near shock waves and turbulent structures away from these discontinuities; (2) proper modelling of the small subgrid scales for supersonic combustion, including effects from compressibility on mixing and combustion; (3) highly detailed kinetic mechanisms (the Warnatz scheme including 9 species and 38 reactions is adopted) accounting for the formation and recombination of radicals to properly predict flame anchoring. Numerical results reveal the complex topology of the flow under investigation. The importance of baroclinic and dilatational effects on mixing and flame anchoring is evidenced. Moreover, their effects on turbulence-scale generation and the scaling law are analysed.     

A computational study of supersonic combustion behind a wedge-shaped flameholder

In this study, large eddy simulation (LES) has been used to examine supersonic flow, mixing, self-ignition and combustion in a model scramjet combustor and has been compared against the experimental data. The LES model is based on an unstructured finite-volume discretization, using monotonicity-preserving flux reconstruction of the filtered mass, momentum, species and energy equations. Both a two-step and a seven-step hydrogen–air mechanism are used to describe the chemical reactions. Additional comparisons are made with results from a previously presented flamelet model. The subgrid flow terms are modeled using a mixed model, whereas the subgrid turbulence–chemistry interaction terms are modeled using the partially stirred reactor model. Simulations are carried out on a scramjet model experimentally studied at Deutsches Zentrum für Luft- und Raumfahrt consisting of a one-sided divergent channel with a wedge-shaped flame holder at the base of which hydrogen is injected. The LES predictions are compared with experimental data for velocity, temperature, wall pressure at different cross sections as well as schlieren images, showing good agreement for both first- and second-order statistics. In addition, the LES results are used to illustrate and explain the intrinsic flow, and mixing and combustion features of this combustor.     

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.      

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.     

Experimental investigation of combustion mechanisms of kerosene-fueled scramjet engines with double-cavity flameholders

A scramjet combustor with double cavitybased flameholders was experimentally studied in a directconnected test bed with the inflow conditions of M = 2.64,Pt = 1.84 MPa,Tt = 1 300 K.Successful ignition and selfsustained combustion with room temperature kerosene was achieved using pilot hydrogen,and kerosene was vertically injected into the combustor through 4×φ 0.5 mm holes mounted on the wall.For different equivalence ratios and different injection schemes with both tandem cavities and parallel cavities,flow fields were obtained and compared using a high speed camera and a Schlieren system.Results revealed that the combustor inside the flow field was greatly influenced by the cavity installation scheme,cavities in tandem easily to form a single side flame distribution,and cavities in parallel are more likely to form a joint flame,forming a choked combustion mode.The supersonic combustion flame was a kind of diffusion flame and there were two kinds of combustion modes.In the unchoked combustion mode,both subsonic and supersonic combustion regions existed.While in the choked mode,the combustion region was fully subsonic with strong shock propagating upstream.Results also showed that there was a balance point between the boundary separation and shock enhanced combustion,depending on the intensity of heat release.     

Detonation combustion of hydrogen in a Laval nozzle under rarefied atmosphere conditions

Detonation combustion of a hydrogen-air mixture entering an axisymmetric convergent-divergent nozzle at a supersonic velocity is considered under atmospheric conditions at altitudes up to 24 km. The investigation is carried out on the basis of the two-dimensional gasdynamic Euler equations for a multicomponent reacting gas. The limiting altitude ensuring detonation combustion in a Laval nozzle of given geometry is numerically established for freestream Mach numbers 6 and 7. The possibility of the laser initiation of detonation in a supersonic flow of a stoichiometric, preliminarily heated hydrogen-air mixture is experimentally studied. The investigation is carried out in a shock tube under conditions simulating a supersonic flow in the nozzle throat region.     

双燃式超燃发动机冷态内流场的数值研究

研究了双燃式一体化通道(包含进气道、双燃式燃烧室和尾喷管)的冷态内流场特性.首次在激波风洞中对内流场进行纹影照相,用TVD格式求解三维全N-S方程对喷管和一体化通道进行分区数值模拟,并考察了几何参数对内流场的影响.结果表明对典型工况(h     

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。     

高马赫数下超声速燃烧的自点火查表方法

超燃冲压发动机燃烧室工作在高马赫数工况时, 入口来流空气的总焓非常高, 自点火在高焓条件下成为维持火焰稳定的重要物理化学过程. 本文借鉴火焰面/进度变量模型的降维思路, 发展了一种基于化学动力学的自点火建表方法. 通过定义混合分数和进度变量将复杂多维的化学反应降维, 并成功将数据库方法结合到现有的大涡模拟求解器中. 经过测试和验证, 该方法初步具备对超声速自点火燃烧进行仿真描述的能力. 针对自点火诱导的超声速燃烧问题开展数值模拟, 该方法通过查表的方式有效降低了化学反应求解过程中的计算量. 在采用详细化学反应机理时能够准确地再现自点火行为和火焰结构, 并且预测的温度和重要组分分布与实验吻合较好.      

双模态发动机的模态鉴别方法

双模态冲压发动机的不同燃烧模态具有不同的稳焰机制和流态特征,并且在模态转换时伴随着显著的推力变化. 因此,准确判断燃烧模态,对于捕捉发动机的燃烧区位置/范围、释热分布特征,以及为进一步优化燃烧室的设计(流道结构和供油布局) 具有重要意义. 目前尚无鉴别模态的有效试验方法,本文提出了一种模态鉴别的试验方法,并在超燃直连台上开展验证试验. 试验中使用的测量技术包括:壁面静压、高速阴影/纹影、多通道可调谐二极管吸收光谱和高能态碳氢自由基CH* 自发光成像. 利用多种测量方法的组合,可以同时获得燃烧室中气流静温、速度、马赫数分布,释热分布以及燃烧区位置/范围. 这些试验数据能够用于判别模态,并获得不同模态的流动和火焰特征.      

飞行Ma12条件超燃发动机流场及燃烧特征分析

为提升针对高马赫数发动机的模拟能力,对计算方法进行了可压缩性修正,并针对飞行Ma12条件下超燃冲压发动机进行了多状态三维数值模拟,分析了发动机内波系、参数以及燃烧性能特征.研究结果表明:(1)修正后的方法计算所得激波位置及强度与试验值吻合,在激波串模拟、高马赫数发动机模拟上均展现了更优的能力.(2)发动机内形成激波与反...     

High order numerical simulation of the thermal load on a lobed strut injector for scramjet applications

In this paper, the thermal load on an actively cooled lobed strut injector for scramjet (supersonic combustion ramjet) applications is investigated numerically. This requires coupled simulations of the strut internal and external flow fields together with the heat conduction in the solid injector body. In order to achieve a fast mixing, the lobed strut is positioned at the channel axis to inject hydrogen into the core of a Mach 3 air stream. There it is exposed to the extremely high temperatures of the high speed flow. While the external air and hydrogen flows are supersonic, the strut internal hydrogen flow is mainly subsonic, in some regions at very low Mach numbers. To enable a simulation of the internal flow field which ranges from very low to very high Mach numbers (approximately Mach 2.25 at the nozzle exit), a preconditioning technique is employed. The compressible finite‐volume scheme uses a spatially fourth order multi‐dimensional limiting process discretization, which is used here for a first time to simulate a geometrically and fluid mechanically highly complex problem. It will be demonstrated that besides its high accuracy the multi‐dimensional limiting process scheme is numerically stable even in case of demanding practical applications. The coupled simulation of the lobed strut injector delivers unique insight into the flow phenomena inside and outside the strut, the heat fluxes, the temperature distribution in the solid material, the required hydrogen mass flux with respect to cooling requirements and details concerning the conditions at the exit of the injector. Copyright © 2016 John Wiley & Sons, Ltd.     

Initiation and stabilization of combustion at low pressure and temperature by means of a nonequilibrium electric discharge

The pattern of supersonic (M = 2) flow near the surface of a plate at points on which propane jets are injected normal to the air flow direction is investigated. For the initiation and intensification of the chemical reactions a nonequilibrium discharge is used. This is created between an anode oriented along the flow, the plate surface, and a metal interceptor mounted on the plate. The results of a schlieren visualization of the flows developed are presented. Spectroscopic studies show that the distribution of plasmochemical reaction products has a number of fundamental differences as compared with the case of propane injection along the plate surface. A comparative analysis of these distributions for identical gasdynamical experimental conditions is important for testing calculation models of reactive-mixture supersonic flows in which electric discharges are used for ignition and the stabilization of combustion.     

高马赫数超燃冲压发动机技术研究进展

吸气式高超声速飞行在空间运输和国家空天安全领域具有极高价值,超燃冲压发动机是其核心动力装置.目前飞行马赫数4.0～7.0超燃冲压发动机技术日趋成熟,发展更高速的飞行动力技术成为今后临近空间竞争焦点之一.本文对飞行马赫数8.0～10.0的高马赫数超燃冲压发动机技术进行了分析和综述.首先论述其亟待解决的关键问题和技术,分别包括高焓离解与热化学非平衡效应、超高速气流燃料增混与燃烧强化技术、高超声速燃烧与进气压缩的匹配及工作模态、高焓低雷诺数边界层流动及其控制方法、高焓低密度流动/燃烧的热防护技术,以及高马赫数发动机的地面试验风洞技术.然后,进一步介绍了国内外高焓激波风洞与驱动技术以及国内外典型的地面和飞行试验进展.进而针对推进和热防护的总体性能评估、高马赫数发动机内凸显的高焓离解与热化学非平衡效应、超高速气流燃料增混和燃烧强化技术综述了相关研究进展及结论,讨论了高马赫数超燃冲压发动机的可行性以及各关键技术的特点.最后进行了总结并对后续研究提出了几点建议.     

Numerical analysis of supersonic combustion ramjet with upstream fuel injection

This paper describes possible fuel injection scheme for airbreathing engines that use hydrocarbon fuels. The basic idea is to inject fuel at the spike tip of the supersonic inlet to achieve mixing and combustion efficiency with a limited length combustion chamber. A numerical code, able to solve the full Navier–Stokes equations in turbulent and reacting flows, is employed to obtain numerical simulations of the thermo‐fluidynamic fields at different scramjet flight conditions, at Mach numbers of M=6.5 and 8. The feasibility of the idea of the upstream injection is checked for a simple axisymmetric configuration and relatively small size. The results are discussed in connection with the potential benefits deriving from the use of new ultra high temperature ceramics (UHTC). Copyright © 2003 John Wiley & Sons, Ltd.     

乙烯燃料超燃燃烧室流动特性与燃烧稳定性研究

在低飞行马赫数条件下,乙烯燃料超燃冲压发动机为实现成功点火及稳定燃烧,常使用先锋氢引燃乙烯,本文通过试验研究了多种喷注方案下的超燃燃烧室流动特性、火焰传播特性及燃烧稳定性,喷注方案包括单先锋氢、单乙烯和组合喷注方式.超燃燃烧室入口马赫数为2.0,总温为953 K,总压为0.82 MPa.多种非接触光学测量手段被应用于超...     

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.