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.      

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。     

基于液体碳氢燃料的旋转爆轰燃烧特性研究

为了探索液体碳氢燃料参与旋转爆轰所产生的不完全燃烧现象,采用守恒元与求解元方法,开展柱坐标系下的汽油/空气两相旋转爆轰燃烧室三维数值模拟研究,针对燃料喷注压力和反应物当量比对旋转爆轰流场结构及燃烧室性能的影响进行分析。分析结果表明:保持总当量比为1.00,随着燃料喷注压力的上升,燃烧室内燃料不均匀分布增强,产生局部富燃区,燃料在燃烧室未能完全反应,导致燃烧室燃料比冲下降;保持喷注压力不变,减小当量比,在贫燃工况下依然存在局部富燃区,导致燃烧室内出现不完全燃烧现象,降低燃烧室比冲性能。由此可知,反应物喷注方案对气液两相旋转爆轰的不完全燃烧有显著影响。     

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。     

基于动态分区概念的高超声速燃烧大涡模拟

本文基于动态分区概念开展了亿级网格的高马赫数全尺寸超燃冲压发动机内外流耦合一体化改进延迟分离涡(IDDES)模拟研究. 研究建立了包括动态分区火焰面湍流燃烧模型(DZFM)、分区自适应化学(Z-DAC)和分区并行自适应建表(Z-ISAT)的完整动态分区燃烧模拟框架, 并通过1.15亿网格的马赫数12 REST标准高超声速燃烧室模型初步验证了分区模拟框架的保真性. DZFM通过分区解耦的思想既准确表征了当地湍流化学交互作用关系, 又有效提升了整场湍流燃烧的计算效率. Z-DAC和Z-ISAT通过在分区框架内对化学反应机理进行动态实时简化和建表查询, 可进一步提升当前分区内化学反应的求解效率. 基于1.25和1.4亿网格动态分区框架对比分析了马赫数10条件下中心支板(strut)和壁面撑挡型(pylon)两类构型氢气高超声速燃烧室特性. 支板或撑挡结构均诱发了明显的边界层分离和头部回流区, 由此两种燃烧室均出现了较长区域的喷注点前部燃烧现象. 基于Borghi图的数值分析表明当前氢气高超声速燃烧室中广泛存在扩散控制为主的火焰面模式, 效率提升的瓶颈在于高效增混. 壁面撑挡燃烧室具有较高的穿透深度和近场混合效率, 因而燃烧效率高于净推力准则80%, 相应的比冲1234 s也远高于中心支板燃烧室的437 s. 分区自适应化学方法在将近一半的计算域上降低了反应求解计算代价, 特别是在无燃料区反应机理的简化幅度更加明显. 相比与传统的有限速率PaSR模型, DZFM模型实现了高达11倍的加速比.      

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.      

超燃冲压发动机燃烧流场及火焰稳定初步研究

采用壁面燃料喷射并结合凹槽设施作为火焰稳定器是超燃冲压发动机设计的理想方案,本文采用非定常数值模拟研究了带凹槽的超燃冲压发动机壁面横向喷射乙烯的火焰稳定过程。结果表明:在燃烧室入口马赫数2、静温530K、静压0.1MPa条件下,冷流流场达到稳定所需时间约为2ms;当凹槽内喷油当量比为0.1时,火焰稳定模式为燃料尾迹和凹槽共同形成的回流区稳定模式;当凹槽内喷油当量比为0.315时,火焰稳定模式完全处于凹槽回流区稳定模式;当凹槽前端壁面注油当量比为0.05时,火焰稳定模式为凹槽回流区稳焰模式;当凹槽前端壁面注油当量比为0.2时,火焰稳定模式为射流回流区和凹槽回流区稳焰模式。     

提高高马赫数超燃冲压发动机推力的理论方法

斜爆轰发动机和激波诱导燃烧冲压发动机在高马赫数吸气式发动机中具有重要应用前景,但是斜爆轰发动机是否具有足够大的净推力,还是一个未知的问题,因此需要对高马赫数冲压发动机的推进性能以及提高推力的方法进行理论研究.本文主要分为3部分.第1部分理论研究了超燃冲压发动机中的爆燃波和爆轰波的传播特性.保证发动机稳定燃烧是提高推力的...     

超燃燃烧室气流参数诊断

基于可调谐二极管激光器吸收光谱技术, 利用7185.597cm^{-1}, 7444.35cm^{-1} + 7444.37cm^{-1}(重合吸收线)两条H_2O吸收线, 采用分时扫描-直接探测策略组建多光路吸收测量系统, 在4kHz的测量频率下, 定量测量了燃烧室气流的静温、水蒸气浓度和流向速度. 利用位移机构, 在以C_2H_4为燃料的超燃直连式试验台中, 在单次试验中同时诊断燃烧室内某截面和燃烧室出口的多气流参数的截面分布. 利用燃烧室出口截面的水蒸气浓度分布, 并结合壁面静压计算燃烧效率; 利用燃烧室出口截面的静温和速度分布, 获得出口气流马赫数分布; 利用凹腔后部某截面的温度和水蒸气浓度分布, 判读了凹腔附近流场特征.      

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.     

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.     

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.     

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.     

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.     

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.     

模型超燃冲压发动机内着火过程分析

在燃烧室入口来流为Ma=2.64、T0=1483K、P0＝1.65MPa、T=724K、P=76.3kPa条件下,采用高速摄影和连续激光高速纹影对等截面型开窗燃烧室内氢气射流自燃过程、火花塞点燃氢气过程和引导氢气火焰点燃煤油过程进行了观测,获得了燃烧室内着火过程中火焰和流场波系结构的动态演化过程;观察到了初始火焰区首先起始于燃烧室下游,并逆流传播实现发动机着火的过程;分析表明燃料能否着火、以及着火位置与燃料着火时间、燃烧室流速和火焰稳定器安装情况相关,多火焰稳定区延长了燃料驻留时间,使燃料更容易着火。关键词 超燃冲压发动机,点火过程,火焰传播,火焰稳定器      

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...