RBCC直扩燃烧室煤油喷雾燃烧火焰稳定与放热规律的数值模拟

扩张构型燃烧室的燃烧流动细节与放热规律是RBCC发动机设计中的核心技 术。采用湍流流动的分离涡（DES）计算方法,数值计算了RBCC燃烧室以凹腔作为火焰 稳定器的液态煤油喷雾燃烧三维两相流动。针对逐级扩张的RBCC燃烧室构型,详细研究了不 同来流状态下的喷雾燃烧流动特征以及液态煤油分级喷注的放热规律。研究表明,高来流总 温条件下,凹腔火焰稳定器可起到驻留火焰的作用,在相对较低来流总温条件下,凹腔并非 是实现火焰稳定的充分条件,必须采用其他方式补偿液态燃料蒸发吸热所损失的热量。考虑 到扩张构型的几何通道承受的压力提升范围有限,燃料喷注位置不宜安置在燃烧室上游流场 ;为了实现最大的燃烧效率以及发动机推力,采用前后级辅助喷注的方式是目前可行的解决 措施。     

两种优化组合式燃料喷注方案的凹腔稳焰特性实验研究

在马赫数2.0,总压0.98 MPa和总温920 K的超声速来流条件下,针对现有常见的凹腔组合式燃料喷注方案出现的燃烧不稳定和火焰吹熄现象,通过改变凹腔上游壁面双路燃料喷注的位置,设计了两种优化的凹腔组合式喷注方案,并对不同燃料喷注方案下的火焰稳定过程进行研究。通过高速摄影和CH*基自发辐射成像技术,详细观测了后缘突扩凹腔燃烧室中乙烯火焰传播过程。研究表明,原始的喷注方案容易发生火焰振荡,并伴随着火焰回传现象以燃烧模式的转换;当量比超过0.3时,就难以实现稳定燃烧,并出现火焰吹熄现象。两种改进的喷注方案均能增强燃料射流与凹腔的相互作用,可在更宽燃料喷注当量比范围内维持火焰不被吹熄。相比于增加上游喷注与凹腔前缘距离的喷注方案而言,增加双路燃料喷注之间距离的喷注方案的稳焰效果更好,燃烧反应区也更加靠近凹腔前缘,燃烧释热也更强。这种喷注方案可为超燃冲压发动机燃烧室中凹腔燃料喷注方案的优化设计提供参考。     

当量比对超声速燃烧室性能影响的数值研究

采用欧拉-拉格朗日法在来流Ma=2的条件下,对带支板凹腔组合结构的煤油超燃燃烧室的内流场进行数值计算,分析了燃烧室下游支板不同当量比对燃烧室燃烧流场的影响,并对燃烧室的性能做了定量分析。研究表明,随下游支板燃料当量比增加,燃烧反压对燃烧室上游影响加重,流动分离区扩大,上游燃料发生亚声速燃烧状态,且亚声速燃烧区域变大。在支板和凹腔共同作用下,凹腔后方形成了亚声速燃烧区和超声速燃烧区,当量比增加时超声速燃烧区减小,亚声速燃烧区扩大,从而有利于燃料的充分混合和燃烧。随当量比增加,燃烧室总压恢复系数和推力增加,燃料消耗率和比冲量减小。     

富氢/富氧燃气温度对同轴直流气-气喷嘴性能的影响

通过求解使用k-ε湍流模型的Navier-Stokes方程组对采用同轴直流气-气单喷嘴燃烧室的燃烧流场进行数值模拟,对比分析了富氢/富氧燃气推进剂与常温氢气/氧气推进剂条件下的燃烧流场、燃烧室室壁和喷注面板处的燃气温度,研究了富氢/富氧燃气温度变化对燃烧流场和燃烧室热载的影响。数值结果表明:富氢/富氧燃气气-气喷嘴的燃烧性能较好,但热载较高;富氢/富氧燃气温度一定范围内提高对燃烧性能影响不明显,而热载增加。     

轴对称结构RBCC发动机超燃模态性能分析

为了改善RBCC发动机超燃模态的性能,设计了轴对称结构燃烧室结合小支板组进行燃料喷注的发动机构型.通过煤油的3步简化动力学模型,对不同燃料喷注方式下的发动机性能进行计算分析.结果表明,基于本设计的发动机,让支板火箭工作于小流量富燃状态,可实现超燃模态的可靠点火和稳定燃烧；采用一级支板结合二级壁喷的燃料喷注方式,可获得相...     

凹腔布局对高超声速飞行器气动-推进性能影响

采用二维耦合隐式N-S方程和标准k-ε湍流模型,对高超声速飞行器在发动机通流状态下的内外流场进行了数值仿真,研究了超燃冲压发动机燃烧室中单凹腔和双凹腔串并联布局对飞行器气动-推进性能的影响。发现因粘性而产生的摩阻、摩升以及摩擦力引起的俯仰力矩较压阻、压升以及压力引起的俯仰力矩很小,对于飞行器整体性能而言,可忽略;凹腔之间距离的长短对飞行器气动-推进性能影响强烈,短距离凹腔并联使得燃烧室主流压力抬高得更大,短距离凹腔串联使得上游凹腔对下游凹腔流场影响更大;同时,性能高的凹腔组合在一起能显著提高飞行器整体性能。     

气氢/液氧同轴喷注单元细节对燃烧特性的影响

为了研究气氢/液氧同轴直流式喷嘴的结构参数细节对燃烧特性的影响,对单喷嘴燃烧室的燃烧流场进行了数值模拟.重点研究了氧喷嘴缩进深度、氧喷嘴出口壁厚和氢氧喷注速度比3个参数对燃烧效率和稳定性的影响规律.研究表明：上述喷嘴结构参数细节是影响气氢/液氧同轴直流式喷嘴燃烧特性的重要因素,其中适当提高氧喷嘴缩进深度或氢氧喷注速度比对燃烧效率有显著改善,而适当提高氧喷嘴出口壁厚对燃烧稳定性有显著改善.     

燃料喷注位置对于RBCC超燃模态性能的影响

为实现二元结构火箭基组合循环(RBCC)发动机在超燃模态下较优的工作性能,开展了数值模拟研究。使用二阶TVD格式差分算法,结合十二步乙烯简化动力学模型,分析了RBCC超燃模态下的冷热态流场,评定燃料喷注位置对发动机性能的影响。数值模拟结果表明,支板火箭关闭情况下,乙烯燃料RBCC发动机可在流道内组织燃烧、建立室压;将燃料在支板与凹腔中间靠上游位置喷注,可获得较好的发动机总体性能,此时发动机内推力增益可达9%以上;支板火箭底部的高温低速回流区有助于维持燃料高效燃烧释热。     

基于凹腔火焰稳定器的亚燃冲压发动机点火性能研究

通过碳氢燃料亚燃冲压发动机直连式试验,对凹腔火焰稳定器的点火性能进行了初步研究.试验分别在高能火花塞及氰气引导火焰两种情况下成功实现了可靠点火.结果表明,基于凹腔火焰稳定技术的亚燃冲压发动机的点火性能与燃料喷注压降及喷注方式密切相关.在壁面喷注燃料的方式下,发动机容易实现可靠点火,而在中心喷注燃料的情况下,发动机很难被点燃.此外,试验还发现,发动机的喉部尺寸对采用这类结构的哑燃冲压发动机凹腔内的压力影响较小,因此,喉部尺寸的变化对其点火性能的影响也较小.     

气氢/液氧同轴喷注器变工况燃烧流场相似性

为研究不同室压工况下气氢/液氧燃烧流场的相似性,设计了喷注器试验件,并采用数值仿真和热试验的方法对气氢/液氧喷注器的喷雾燃烧流场进行了研究。数值仿真选取试验件的1/6进行三维稳态计算,其中湍流模型采用SST k-ω模型、化学反应采用考虑氢氧6组分9步反应机理的涡耗散概念模型、液氧液滴采用离散相模型,共进行了2.8～9.8 MPa范围内8种典型工况的数值仿真。热试验采用气氢/液氧推进剂,进行了4.5 MPa、5.4 MPa和6.8 MPa这3种不同室压工况共4次挤压热试验,采用量热式水冷身部对燃烧室壁面热流进行了测量。仿真和试验结果表明：对于气氢/液氧同轴直流喷注器,在混合比、氢氧温度和喷注速度相同的情况下,当室压大于液氧临界压力时的燃烧流场具有相似性;而室压小于液氧临界压力时的燃烧流场与大于临界压力的燃烧流场结构存在差异。     

壁面孔型对超声速气流中喷注特性的影响

超声速气流中,燃料与来流空气的高效混合是燃烧室实现点火、稳焰及高效燃烧组织的前提。国内外研究者已对比研究了不同壁面孔型对超声速气流中喷注、混合特性的影响,相比于最常见的圆形喷孔,菱形、楔形-半圆、箭形及针形等喷孔用于超声速气流燃料喷注时,不仅有利于降低喷孔前缘边界层的分离,而且也有利于提升射流穿透深度;相比于单孔喷注,组合型喷孔能进一步增强燃料与来流空气在射流远场的混合效果。通过综述各型喷孔的喷注特性,分析提出了适用于超声速燃烧组织的壁面喷注孔型及其工程应用条件。     

Numerical investigation of the impact of asymmetric fuel injection on shock train characteristics

Numerical simulations are carried out to investigate the impact of asymmetric fuel injection on shock train characteristics using the commercial-code FLUENT. The asymmetry of fuel injection is examined by changing the fuel flow rates of the upper and lower wall fuel injectors. The numerical approach solves the two-dimensional Reynolds-averaged Navier–Stokes (RANS) equations, supplemented with a k-ω model of turbulence. As a result, different ways of fuel injections will always lead to shock train transitions, with the variations of shock train structure, strength and leading edge position. For symmetric fuel injection, the flowfield of the isolator is quite asymmetric with the boundary layer of the upper wall side developing much stronger than that of the lower wall, which is due to the heterogeneity of the incoming flow. Regarding to asymmetric fuel injection with more of lower wall side, though the pressures in the combustor are nearly the same, the first shock of the shock train converts between ‘Distinct symmetric X type shock’ and ‘Obscure and weaker asymmetric shock’ and the shock train leading edge moves upstream with the increase of the asymmetry level. With regard to asymmetric fuel injection with more of upper wall side, ‘incomplete asymmetric X type shock’ occurs and the shock train structures keep nearly the same with low level of fuel injection asymmetry. Unexpected results like unstart will happen when increasing the level of fuel injection asymmetry. And the isolator will come back to normal state by decreasing the differential of upper and lower wall sides fuel injections.     

Numerical investigation of the nonreacting and reacting flow fields in a transverse gaseous injection channel with different species

The mixing and combustion process has an important impact on the engineering realization of the scramjet engine. The nonreacting and reacting flow fields in a transverse injection channel have been investigated numerically, and the predicted results have been compared with the available experimental data in the open literature, the wall pressure distributions, the separation length, as well as the penetration height. Further, the influences of the molecular weight of the fuel and the jet-to-crossflow pressure ratio on the wall pressure distribution have been studied. The obtained results show that the predicted results show reasonable agreement with the experimental data, and the variable trends of the penetration height and the separation distance are almost the same as those obtained in the experiment. The vapor pressure model is suitable to fit the relationship between the penetration height, the separation distance and the jet-to-crossflow pressure ratio. The combustion process mainly occurs upstream of the injection port, and it makes a great difference to the wall pressure distribution upstream of the injection port, especially when the jet-to-crossflow pressure ratio is large enough, namely 17.72 and 25.15 in the range considered in the current study. For hydrogen, the combustion downstream of the injection port occurs more intensively, and this may be induced by its smaller molecular weight.     

轴对称结构RBCC发动机超燃模态试验和数值模拟

为研究轴对称结构RBCC发动机超燃模态下的点火和燃烧性能,进行了地面直连试验。采用中心支板火箭与小支板组喷注相结合的方式作为点火和火焰稳定方式,并对燃料喷注方案进行了研究。试验与数值模拟结果表明,采用这种点火方式能实现轴对称结构RBCC发动机的可靠点火和稳定燃烧。二次燃料采取多级喷注的方式能充分利用流道中的氧气,实现较充分的燃烧,但应控制燃料喷注比例。双支板组的加入,能促进燃料与中心空气流的充分掺混,提升燃烧效率,获得较优的燃烧性能。     

Numerical investigation on hydrogen combustion in a scramjet with 3D sidewall compression inlet

Numerical simulations were employed to analyze the flowfield of a scramjet with three-dimensional (3D) sidewall compression inlet, and the effect of inlet distortion on the mixing and combustion process was examined. The numerical approach solved the compressible Reynolds Averaged Navier–Stokes (RANS) equations supplemented with a finite rate chemical reacting model for the combustion of hydrogen fuel and air. Turbulence closure was achieved using Menter shear-stress transport (SST) model. To verify the accuracy of the simulation, the computed wall pressure was compared with the experimental data of the direct-connect combustor test. The metrics employed in the simulations included qualitative assessments related to flow structure as well as quantitative values of fuel mixing efficiency, combustion efficiency and static pressure distribution. Intake sidewalls were found to strongly affect the inlet flow structure, which became more complex in the nonuniform flowfield. The shock train system affected the combustion region located upstream of the injection and led to pairs of asymmetric separation bubbles. Nevertheless, the shock train system dissipated due to the reactions, the combustion patterns of each fuel jets in downstream region were nearly identical, and the degree of improvement of mixing and combustion efficiency near the downstream injectors was less than that near the upstream injectors.     

带不同长度凹腔超声速燃烧数值研究

对带不同长深比凹腔的燃烧室三维燃烧流场进行数值模拟,研究了燃烧室流场结构。结果表明：液体碳氢燃料穿透深度较小;凹腔长深比对燃烧效率、总压损失影响较小,对燃烧室阻力影响显著。     

Experimental study of a flush wall scramjet combustor equipped with strut/wall fuel injection

In this study a flush wall scramjet combustor is tested in a supersonic incoming air flow with the Mach number of 3 which is generated by an air vitiation heater producing the stagnation temperature of 1505 K. Using liquid kerosene as the fuel, the flame is stabilized by means of a centrally mounted O₂ pilot strut after being ignited by a plasma torch. During experimental measurements, the fuel is injected with a constant equivalence ratio of 0.8 according to specified strut/wall injection ratios, i.e., a portion of the fuel amount is injected from the strut while the rest is injected from the wall. The strut and wall injectors are arranged at the same axial position. The combustion performance and wall temperature gradients are evaluated with various fuel feeding ratios between the wall and the strut. Experimental results show, when the equivalence ratio is constant and the axial injection position is fixed, the combustion characteristics vary significantly with the strut/wall fuel feeding ratio, especially when this ratio is close to its lowest and highest limits. Among the four fuel feeding ratios examined, the strut only injection mode and the average distributed strut/wall injection mode show the best combustion performance. However, the strut/wall injection mode produces a smaller wall temperature gradient compared to the strut only injection mode, which is due to the significant film cooling effect caused by the wall injected liquid kerosene.     

Large eddy simulation based studies of jet–cavity interactions in a supersonic flow

Interactions of a cavity flameholder with an upstream injected jet in a Ma 2.52 supersonic flow are investigated numerically. A hybrid RANS/LES (Reynolds-Averaged Navier–Stokes/Large Eddy Simulation) method acting as wall-modeled LES is adopted, for which the recycling/rescaling method is introduced to treat the unsteady turbulent inflow. Patterns of the fluid entrainment into the cavity and escape from the cavity are identified using a scalar-tracing method. It is found that the jet–cavity interactions remarkably enhanced the mass exchange between the fluids in and out of the cavity, resulting in reduced residence time of the cavity fluids. Increasing the distance between the fuel injection and the cavity leading edge tends to attenuate the jet–cavity interactions, leading to weaker mass exchange. Raising the injection pressure appears to enhance the jet–cavity interactions, resulting in a shorter residence time of the cavity fluids. Moreover, the mass decay processes for the fuel and air within the cavity are basically the same while the entrainment processes for the fuel and air into the cavity seem quite different.     

双脉冲发动机点火过程三维数值模拟

以喷射棒式双脉冲发动机燃烧室、级间隔离装置和喷管一体化为研究对象,采用数值仿真技术对Ⅱ脉冲点火过程三维流场特性进行分析研究。计算结果表明,点火初期燃气压力波峰超前于火焰峰到达级间隔离装置,并以压强冲击波形式传播,Ⅱ脉冲燃烧室相对高压区位置不断发生改变;级间孔打开过程对药柱末端压强影响较大,但对Ⅱ脉冲燃烧室压强整体上升过程影响较小;级间孔打开后,燃气经级间孔加速后形成高度欠膨胀射流,并在Ⅰ脉冲燃烧室内形成非对称带状低压区;级间孔分布的非对称性,导致压强及温度在发动机燃烧室中呈现显著的三维分布特性;高温区出现在隔板附近,而在装药前端、装药末端及外围级间孔轴线附近出现低温区。     

高温风洞收集口喷水降温数值仿真研究

针对高温风洞中扩压器前段壁面防热问题，提出对高温气流外缘喷水降温的方法。通过在收集器入口与喷管出口间安装喷水环，利用液态水汽化吸热对高温气流进行降温，使扩压器壁面形成低温保护层。为了解该方法降温效果，本文利用DPM、组分输运等模型的耦合建立了超声速两相流CFD模型，对向超声速热气流喷水进行降温的过程进行了数值计算，计算结果表明，扩压器启动后有显著的降温保护效果。同时，为探索风洞排气背压和喷水量对风洞流场和壁面降温效果的影响，通过计算得出了变排气背压、变喷水量与降温效果之间的关系，为高温风洞收集口喷水降温装置的优化设计提供了参考。