斜爆轰发动机流动机理分析

为了研究高Mach数超燃冲压发动机和斜爆轰发动机的内流场燃烧流动机理，首先用CJ爆轰理论对超燃冲压发动机的内流场特性进行了理论分析，给出了燃烧室流场的气动规律，理论分析结果与现有实验结果吻合得非常好.其次，根据理论分析结果，提出了高Mach数超燃冲压发动机和斜爆轰发动机的气动设计原则.最后，根据提出的气动设计原则，设计了高Mach数斜爆轰发动机，飞行Mach数为9，对斜激波诱导燃烧机理开展了二维数值模拟研究.数值模拟结果表明，在高Mach数下，斜爆轰发动机燃烧室内可以得到稳定的燃烧流场.      

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.      

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.      

JF12激波风洞高Mach数超燃冲压发动机实验研究

针对高Mach数（Ma ≥ 7）超燃冲压发动机高气动阻力下的燃烧组织问题,提出一种双突扩燃烧室结构方案.使用数值模拟方法考察了射流与双突扩燃烧室组合方式的混合燃烧特性.设计了双突扩超燃冲压发动机模型,在力学研究所JF12长试验时间激波风洞内,开展了Ma=7.0和Ma=9.5的氢燃料点火和燃烧试验对比.在风洞有效试验时间100 ms内,实现了Ma=7.0和Ma=9.5超燃冲压发动机的成功点火与稳定燃烧.在Ma=7.0情况下,进气道采用三维压缩,燃烧室入口设计Mach数Ma_c=2.5,壁面压力分布实验结果显示燃烧放热靠近燃烧室扩张段上游;在Ma=9.5情况下,进气道采用二维压缩,燃烧室入口设计Mach数Ma_c=3.5,由于燃烧室流动速度特别高,燃烧放热靠近燃烧室扩张段下游.      

超声速气流中的爆震推进理论与研究进展

爆震燃烧近似为等容燃烧,理论上其热循环效率高于基于等压燃烧的爆燃燃烧,在超声速推进系统中具有潜在的应用价值.通过总结超声速气流中的爆震推进理论与研究进展,分析其需要解决的关键科学与技术问题,指导未来高超声速发动机的基础研究.文章重点总结了适用于高超声速飞行的斜爆震发动机、超声速脉冲爆震冲压发动机的基础研究进展.其中对斜爆震发动机的应用模式、相关实验研究思路及方法、数值仿真现状进行了总结分析.对超声速脉冲爆震冲压发动机的基础理论研究现状和目前研究的难点进行了梳理.基于爆震燃烧的超燃冲压发动机具有推进系统自增压、燃烧效率高、推力性能好、推进效率高、燃烧室长度短、结构重量轻等优势,文章总结了该发动机当前的发展进程和最新的研究进展,并对其未来的发展方向以及存在的技术问题进行了分析.      

隔离段激波串流场特征的试验研究进展

高超声速推进技术是国际前沿研究, 其中双模态超燃冲压发动机的发展受到极大关注. 作为超燃冲压发动机的重要部件, 隔离段对发动机的性能和高超声速飞行的实现至关重要, 其中所涉及的流动机理问题也极为复杂. 自从高超声速飞行的概念被提出和论证以来, 相关的理论、试验和仿真研究不断取得进展, 但是对其中的机理问题研究仍有待进一步深入. 本文将从试验研究的角度回顾并综述近年来超燃冲压发动机隔离段的研究进展, 结合精细流动测试技术(Nano-tracer Planar Laser Scattering, NPLS)的发展分析了隔离段流场特征, 包括了激波串流场复杂的三维时空结构特点、湍流特性、非线性迟滞运动、不启动流场特征以及激波前缘检测等. 从风洞设备、隔离段设计、测试技术等方面对隔离段的试验研究进行了分类比较和论述, 对今后隔离段试验研究提出了建议.     

超燃冲压发动机先进性量化评估方法

超燃冲压发动机的评估评价对高超声速飞行技术的发展有重要影响。采用层次分析法并结合超燃冲压发动机的技术特点,利用各性能指标的无量纲化参数,研究建立了量化评估超燃冲压发动机先进性的方法与模型,并通过实际算例对其可行性和合理性进行了说明。该方法将复杂的超燃冲压发动机先进性量化评估问题简化为定量分析和数值模拟的过程,对超燃冲压发动机的评估方法和发展方向具有重要意义。     

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。     

弹头激波诱导燃烧特性的数值模拟

为了研究弹头激波诱导燃烧，基于有限体积的考虑化学反应的Navier-Stokes（N-S）方程，对预混氢气-空气化学恰当比时的燃烧流场进行了数值模拟.时间项基于2阶隐式LU-SGS格式，对流项基于Steger-Warming进行离散，化学反应源项采用对角化隐式处理.首先，研究了网格对燃烧爆轰流场结构的影响，并利用Lehr实验结果验证了计算方法的可靠性；其次，研究了弹头的飞行Mach数（Ma=4.18，5.11，6.46）、弹头直径（D=5，10，15 mm）对燃烧流场稳定性的影响.研究表明:计算网格对氢气-空气爆轰流场结构影响很大；弹头直径一定时，氢气-空气燃烧流场稳定性随着飞行Mach数的增大而增强；弹头飞行Mach数一定时，氢气-空气燃烧流场稳定性随着弹头直径减小而增强.      

一种新型的超燃冲压发动机闭式冷却循环

热防护技术是发展高超音速的关键技术之一.为了降低冷却用燃料需求,本文提出了基于闭式Brayton循环的超燃冲压发动机冷却循环,旨在提出一个缓解超燃冲压发动机热防护压力的新方法、新思路,同时可为高超音速飞行器提供足够的电力供应.文中介绍了冷却循环的组成和工作原理,定义了评价冷却循环性能的参数,在考虑系统内、外不可逆性的基础上,完成了系统的性能分析,分析表明该冷却循环可有效地节约冷却用燃料消耗,大幅提升超燃冲压发动机热防护系统性能.     

Numerical simulation of the operation process and thrust performance of an air-breathing pulse detonation engine in supersonic flight conditions

Multidimensional simulations of the unsteady gasdynamic flow in the duct of an air-breathing pulse detonation engine (ABPDE) operating on propane gas and the flow around it in supersonic flight at Mach numbers M of 3.0 and an altitude of 9.3 and 16 km are performed. It is shown that, at a length and diameter of the duct of 2.12 m and 83 mm, respectively, an ABPDE with an air intake and a nozzle can operate in a cyclic mode at a repetition frequency of 48 Hz, with a rapid deflagration-to-detonation transition (DDT) occurring at a distance of 5–6 combustion chamber diameters. To determine the thrust performance of the ABPDE in flight conditions, a series of working cycles were simulated with consideration given to the external flow around the engine. Calculations showed that the specific impulse of the ABPDE is approximately 1700 s. This value is much higher than the specific impulse typical of ramjet engines operating on conventional combustion (1200–1500 s) and substantially lower than the specific impulse obtained for the atmospheric conditions at sea level at zero flight velocity (∼2500 s).     

Three-dimensional numerical simulation of the characteristics of a ramjet power plant with a continuous-detonation combustor in supersonic flight

Multi-variant three-dimensional numerical simulations demonstrate the feasibility of the continuous- detonation process in an annular combustor of a ramjet power plant operating on hydrogen as fuel and air as oxidant in conditions of flight at a Mach number of M ₀ = 5.0 and an altitude of 20 km. Conceptual schemes of an axisymmetric power plant, 400 mm in external diameter and 1.3 to 1.5 m in length, with a supersonic intake, divergent annular combustor, and outlet nozzle with a frusto-conical central body are proposed. Calculations of the characteristics of the internal and external flows, with consideration given to the finite rate of turbulent-molecular mixing of the fuel mixture components with each other and with the combustion products, as well as the finite rate of chemical reactions and the viscous interaction of the flow with the bounding surfaces, have shown that, in these flight conditions, the engine of such a power plant has the following performance characteristics: the thrust, 10.7 kN; specific thrust, 0.89 (kN s)/kg; specific impulse, 1210 s; and specific fuel?consumption 0.303 kg/(N h). In this case, the combustor can operate with one detonation wave traveling in the annular channel at an average velocity of 1695 m/s, which corresponds to a detonation wave rotation frequency of 1350 Hz. It is shown that, an operating combustor has regions with subsonic flow of detonation products, but the flow is supersonic throughout its outlet section.     

LES of supersonic combustion in a scramjet engine model

In this study, Large Eddy Simulation (LES) has been used to examine supersonic flow and combustion in a model scramjet combustor. The LES model is based on an unstructured finite volume discretization, using total variational diminishing flux reconstruction, of the filtered continuity, momentum, enthalpy, and passive/reactive scalar equations, used to describe the combustion process. The configuration used is similar to the laboratory scramjet at the Institute for Chemical Propulsion of the German Aerospace Center (DLR) and consists of a one-sided divergent channel with a wedge-shaped flameholder at the base of which hydrogen is injected. Here, we investigate supersonic flow with hydrogen injection and supersonic flow with hydrogen injection and combustion. For the purpose of validation, the LES results are compared with experimental data for velocity and temperature at different cross-sections. In addition, qualitative comparisons are also made between predicted and measured shadowgraph images. The LES computations are capable of predicting both the non-reacting and reacting flowfields reasonably well—in particular we notice that the LES model identifies and differentiates between peculiarities of the flowfields found in the experiments.     

Dual/scram-mode combustion limits of ethylene and surrogate endothermically-cracked hydrocarbon fuels at Mach 8 equivalent high-enthalpy conditions

This paper examines the scram/dual-mode combustion limits of hydrocabon fuels within a Mach 8, scramjet combustor. Flight-equivalent flows were delivered to the axisymmetric, cavity combustor via a reflected shock tunnel. Two scramjet fuels were examined: ethylene and a surrogate mixture representing endothermically cracked n-dodecane. Combustion modes were examined via static pressure sensors and through both chemiluminescence imaging, and planar laser induced fluorescence (PLIF) of the OH combustion radical in the combustor exhaust plume. Ethylene-fuelled experiments developed scram-mode combustion under reduced fuelling conditions, experiencing shock wave dominated flowfields. OH PLIF diagnostics indicated such combustion modes developed a ring-like structure of combustion products, primarily axisymmetrically adjacent to the combustor wall. Increased fuelling anchored combustion downstream of the fuel injector, while further increases instigated dual-mode combustion. In this mode, subsonic combustion regions combine with the supersonic coreflow to permit the transfer of information upstream with substantially increased pressure encountered. Optical diagnostics indicate broadly asymmetric, unsteady combustion features. The surrogate mixture representing endothermically cracked n-dodecane experienced rapid onset from no-combustion (optically confirmed) to fully developed dual-mode combustion at critical fuelling rates. OH PLIF signals and chemiluminescence of this fuel were weaker than comparable ethylene cases, indicating potential differences in combustion pathways.     

Thrust characteristics of an airbreathing pulse detonation engine in supersonic flight at various altitudes

The main thrust characteristics, such as thrust force, specific impulse, specific fuel consumption, and specific thrust, of a pulse detonation engine (PDE) with an air intake and nozzle in conditions of flight at a Mach number of 3 and various altitudes (from 8 to 28 km above sea level) are for the first time calculated with consideration given to the physicochemical characteristics of the oxidation and combustion of hydro-carbon fuel (propane), finite time of turbulent flame acceleration, and deflagration-to-detonation transition (DDT). In addition, a parametric analysis of the influence of the operation mode and design parameters of the PDE on its thrust characteristics in flight at a Mach number of 3 and an altitude of 16 km is performed, and the characteristics of engines with direct initiation of detonation and fast deflagration are compared. It is shown that a PDE of this design greatly exceeds an ideal ramjet engine in specific thrust, whereas regarding the specific impulse and specific fuel consumption, it is not inferior to the ideal ramjet.