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.      

超燃冲压发动机推进性能理论分析

超燃冲压发动机的正推力问题和超声速燃烧的稳定性问题是制约超燃冲压发动机发展的两个关键气动物理问题.虽然经过50多年的研究,但是目前国内外对这两个关键问题的机理还没有研究清楚.文章首次将CJ爆轰理论应用于超燃冲压发动机推进性能分析,给出了这两个关键气动问题的理论分析结果.分析结果表明,燃烧室入口空气静温对发动机的推进性能产生重要影响.当爆轰波的爆速大于隔离段内空气来流的速度时,会向隔离段上游传播,导致发动机不起动.飞行Mach数Ma=6~8是超燃发动机的临界不稳定范围,飞行Mach数Ma>9,超声速燃烧将变得稳定.      

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.      

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.      

FD-21风洞Ma=10高超声速推进试验技术探索

针对Mach数8以上（Ma>8）冲压发动机地面试验能力不足问题,基于FD-21高能脉冲风洞,开展了吸气式推进试验技术探索,提升了FD-21风洞的重活塞驱动能力,获得了总压18.66 MPa、总温3 950 K、Ma=9.62、静压436.6 Pa、速度3 km/s的高焓大动压模拟流场,同时发展了高时间分辨率吸收光谱测量技术和基于重模型自由飞原理的发动机推阻测量方法.在此基础上,设计了弯曲激波压缩二元发动机,构建了燃料在线供应与喷注控制、模型悬挂与瞬态释放及相关测量一体的试验系统,在所建立的Ma=9.62风洞模拟环境中进行了集成验证试验,定量测得了有/无氢气射流与空气/氮气超声速气流作用下二元发动机的壁面压力、吸收光谱峰值吸收率、轴向力等数据,并利用纹影观测到了进气道唇口与燃烧室部位的波系特征.多次试验所得的壁面压力、峰值吸收率、轴向力随时间变化曲线均存在2 ms以上的平台,表明二元发动机建立了准定常流动.冷热态及氮气对照组对应的壁面压力分布、峰值吸收率、轴向力等数据呈现出了明显不同,且二者规律近似一致,一方面说明所建立的模拟流场、燃烧诊断技术、发动机推阻测量技术是有效的,另一方面也表明二元发动机实现了点火燃烧、获得有效热功转换,为后续相关研究奠定了良好的基础.      

基于CARS技术的超燃冲压发动机点火过程温度测量

相干anti-Stokes Raman散射（coherent anti-Stokes Raman scattering,CARS）技术作为一种非接触测量手段,已广泛应用于多种发动机模型燃烧室温度测量及地面试验.然而,目前的工作主要集中在稳态燃烧场温度的测量,缺乏用高分辨率的单脉冲来测量瞬变的燃烧火焰温度及组分浓度的研究.基于CARS理论,结合多参数拟合算法,开发了基于MATLAB的CARS光谱计算和拟合程序CARSCF;利用McKenna平面火焰炉在不同工况下进行了温度测量,并与DLR测量结果进行对比,结果显示开发的CARSCF具有较高的测量重复性和准确性;最后将CARS技术应用于测量超燃冲压发动机点火过程中的温度测量,获取了点火过程中的温度.结果显示,在来流Mach数为3的条件下,H₂/air点火过程中温度呈现急剧上升然后缓慢下降,而CARS信号则呈现急剧上升然后急剧下降随后又缓慢上升的趋势,并且在点火过程中最高温度为1 511 K.      

吸气式高超声速飞行器机体/推进一体化内外流非定常耦合方法

基于吸气式高超声速飞行器机体/推进一体化的气动布局设计方式,文章提出了一种内外流一体化流场的耦合求解方法,其中燃烧室内流场采用考虑有限速率化学反应动力学模型的一维非稳态方法求解,进气道和尾喷管外流场采用二维CFD软件计算,进气道与燃烧室在耦合界面处通过一维平均方法实现静温、静压和Mach数等参数传递.并分别以日本国家航空与航天实验室（NAL）的氢燃料燃烧室模型作为内流场验证算例,以某典型高超声速飞行器一体化模型作为内外耦合流场验证算例.研究结果表明:有限速率化学反应准一维方法能较为准确地模拟燃烧室内燃烧流场,提出的内外流场耦合方法能够有效地计算出内外流耦合效应,计算后体压力分布与理论值较接近.该方法可为超燃冲压发动机的性能快速分析和吸气式高超声速飞行器机体/推进一体化的初步分析设计提供重要参考.      

超声速气流中液体横向射流雾化过程数值模拟

为了更加深入了解超燃冲压发动机燃烧室中的燃料雾化机理,对来流Mach数为1.94的超声速气流中液体横向射流的雾化过程进行了数值模拟研究.计算采用Euler-Lagrange方法,液滴二次破碎模型采用K-H/R-T模型.计算结果表明:考虑液滴二次破碎时,采用雾化锥模型获得的射流穿透深度以及液滴速度分布与实验结果符合得很好...     

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。     

超燃冲压发动机喷管推力性能理论预测

超燃冲压发动机发展60多年来,虽然取得了很大的进步,但是对其推力大小的理论评估是一个没有很好解决的问题.超燃冲压发动机的推力主要由喷管产生,因此重点研究了喷管的推力特性.将燃烧室出口参数作为喷管入口边界条件,利用等熵膨胀理论,通过对喷管壁面压力积分,得到了简化的无量纲推力公式,获得了影响推力大小的关键参数和物理规律.理...     

Numerical investigation of supersonic flow breakdown at the inlet duct throttling

The work presents the results of investigating the process of supersonic flow deceleration in a duct of the two-dimensional inlet throttled by variation of the outlet cross-sectional area. An inlet with three external compression shock waves designed for the freestream Mach number M_d = 7 was considered as an example for the investigation. A one-dimensional analysis of the conditions for realization of the supersonic flow deceleration regimes in the inlet duct with two throats — in the inlet entrance and at the inlet duct outlet, has been carried out. The parametric numerical computations of two-dimensional inviscid or turbulent flows in the inlet were performed with the use of the Euler and Navier—Stokes codes of the program package FLUENT. The critical conditions for the nonuniform flow in the outlet throat bringing to choking the inlet duct were determined.     

Application and theoretical analysis of the flamelet model for supersonic turbulent combustion flows in the scramjet engine

In the framework of Reynolds-averaged Navier–Stokes simulation, supersonic turbulent combustion flows at the German Aerospace Centre (DLR) combustor and Japan Aerospace Exploration Agency (JAXA) integrated scramjet engine are numerically simulated using the flamelet model. Based on the DLR combustor case, theoretical analysis and numerical experiments conclude that: the finite rate model only implicitly considers the large-scale turbulent effect and, due to the lack of the small-scale non-equilibrium effect, it would overshoot the peak temperature compared to the flamelet model in general. Furthermore, high-Mach-number compressibility affects the flamelet model mainly through two ways: the spatial pressure variation and the static enthalpy variation due to the kinetic energy. In the flamelet library, the mass fractions of the intermediate species, e.g. OH, are more sensible to the above two effects than the main species such as H₂O. Additionally, in the combustion flowfield where the pressure is larger than the value adopted in the generation of the flamelet library or the conversion from the static enthalpy to the kinetic energy occurs, the temperature obtained by the flamelet model without taking compressibility effects into account would be undershot, and vice versa. The static enthalpy variation effect has only little influence on the temperature simulation of the flamelet model, while the effect of the spatial pressure variation may cause relatively large errors. From the JAXA case, it is found that the flamelet model cannot in general be used for an integrated scramjet engine. The existence of the inlet together with the transverse injection scheme could cause large spatial variations of pressure, so the pressure value adopted for the generation of a flamelet library should be fine-tuned according to a pre-simulation of pure mixing.     

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%.     

超声速混合层MHz级超高频流动可视化实验

研究发展了超高频基于纳米示踪的平面激光散射（nano-tracer planar laser scattering,NPLS）技术。基于多腔并联脉冲激光器技术、棱锥分光与短曝光相机集成技术以及高精度同步控制技术,实现了MHz级流场可视化和精细测量。采用超高频NPLS技术研究了对流Mach数Ma_c=0.17,0.26混合层流场,获得了时间序列的混合层高分辨率NPLS图像。采用阵列型涡流发生器开展流动控制研究,分析涡流发生器对混合层发展的影响特性。通过选取典型涡结构,分析了超声速混合层不同发展阶段的涡运动和发展演化规律。发现混合层中段的不稳定性发展阶段,涡结构以平移和旋转为主,伴随一定的拉伸;混合层后段以变形和破碎为主,有大量小尺度结构产生。并且小尺度结构会受到剪切、大尺度结构以及小激波的影响,发生明显的非定常运动。      

Gasdynamic design of a two-dimensional supersonic inlet with the increased flow rate factor

A two-dimensional inlet of external compression with the increased flow rate factor at high supersonic velocities is constructed by the method of gasdynamic design. Its feature is that a flow with the initial oblique shock wave and the subsequent centered isentropic compression wave is formed over the external compression ramp of the inlet. These waves interact with one another so that a resulting stronger oblique shock wave and a velocity discontinuity arise in front of the entrance to the inlet internal duct. An example of an inlet configuration with the design flow regime corresponding to the Mach number M_d = 7 is considered. The characteristics of this inlet were obtained in the range of the free-stream Mach numbers M = 4–7 with the use of a Navier—Stokes code for turbulent flow. They are compared with characteristics of an equivalent conventional shocked inlet. As computations have shown, the inlet with the isentropic compression wave has much higher values of flow rate factor φ at Mach numbers M < M_d. So, for example, at M = 4 the value φ ≈ 0.72 for it is by 33 % higher in comparison with φ ≈ 0.54 for the equivalent shocked inlet.     

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.