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.     

Multi-objective design optimization of the transverse gaseous jet in supersonic flows

The mixing process between the injectant and the supersonic crossflow is one of the important issues for the design of the scramjet engine, and the efficiency mixing has a great impact on the improvement of the combustion efficiency. A hovering vortex is formed between the separation region and the barrel shock wave, and this may be induced by the large negative density gradient. The separation region provides a good mixing area for the injectant and the subsonic boundary layer. In the current study, the transverse injection flow field with a freestream Mach number of 3.5 has been optimized by the non-dominated sorting genetic algorithm (NSGA II) coupled with the Kriging surrogate model; and the variance analysis method and the extreme difference analysis method have been employed to evaluate the values of the objective functions. The obtained results show that the jet-to-crossflow pressure ratio is the most important design variable for the transverse injection flow field, and the injectant molecular weight and the slot width should be considered for the mixing process between the injectant and the supersonic crossflow. There exists an optimal penetration height for the mixing efficiency, and its value is about 14.3 mm in the range considered in the current study. The larger penetration height provides a larger total pressure loss, and there must be a tradeoff between these two objection functions. In addition, this study demonstrates that the multi-objective design optimization method with the data mining technique can be used efficiently to explore the relationship between the design variables and the objective functions.     

Numerical validation and parametric investigation on the cold flow field of a typical cavity-based scramjet combustor

The three-dimensional coupled implicit Reynolds Averaged Navier–Stokes (RANS) equations and the two equation standard k–ε turbulence model has been employed to numerically simulate the cold flow field in a typical cavity-based scramjet combustor. The numerical results show reasonable agreement with the schlieren photograph and the pressure distribution available in the open literature. The pressure distribution after the first pressure rise is under-predicted. There are five shock waves existing in the cold flow field of the referenced combustor. The first and second pressure rises on the upper wall of the combustor are predicted accurately with the medium grid. The other three shock waves occur in the core flow of the combustor. The location of the pressure rise due to these three shock waves could not be predicted accurately due to the presence of recirculation zone downstream of the small step. Further, the effect of length-to-depth ratio of the cavity and the back pressure on the wave structure in the combustor has been investigated. The obtained results show that there is an optimal length-to-depth ratio for the cavity to restrict the movement of the shock wave train in the flow field of the scramjet combustor. The low velocity region in the cavity affects the downstream flow field for low back pressure. The intensity of the shock wave generated at the exit of the isolator depends on the back pressure at the exit of the combustor and this in turn affects the pressure distribution on the upper wall of the combustor.     

超燃冲压发动机燃烧室流动问题的数值验证

基于三维可压缩RANS方程,结合有限速率化学反应模型,发展了一套模拟多组元化 学 反应与湍流流动的计算程序。通过对超燃冲压发动机燃烧室内的几种典型流动问题,如横 向喷流问题、后向台阶流动问题和氢／空气化学反应问题进行数值模拟,验证计算程序的可 靠性,将数值模拟结果与实验结果进行了对比,结果吻合较好。     

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。     

超声速横向气流中喷雾的数值模拟

对超声速横向气流中的喷雾过程进行了数值模拟,采用二维N-S方程计算气相,应用一次雾化模型和二次雾化模型模拟了喷雾雾化过程,并与实验测量结果进行了对比。研究了湍流度和附面层厚度对液雾穿透深度的影响,发现湍流度和附面层厚度并不是主要的影响因素,认为雾化模型是影响液雾穿透深度的关键因素。     

Performance evaluation and parametric analysis on cantilevered ramp injector in supersonic flows

The cantilevered ramp injector is one of the most promising candidates for the mixing enhancement between the fuel and the supersonic air, and its parametric analysis has drawn an increasing attention of researchers. The flow field characteristics and the drag force of the cantilevered ramp injector in the supersonic flow with the freestream Mach number 2.0 have been investigated numerically, and the predicted injectant mole fraction and static pressure profiles have been compared with the available experimental data in the open literature. At the same time, the grid independency analysis has been performed by using the coarse, the moderate and the refined grid scales, and the influence of the turbulence model on the flow field of the cantilevered ramp injector has been carried on as well. Further, the effects of the swept angle, the ramp angle and the length of the step on the performance of the cantilevered ramp injector have been discussed subsequently. The obtained results show that the grid scale has only a slight impact on the flow field of the cantilevered ramp injector except in the region near the fuel injector, and the predicted results show reasonable agreement with the experimental data. Additionally, the turbulence model makes a slight difference to the numerical results, and the results obtained by the RNG k−ε and SST k−ω turbulence models are almost the same. The swept angle and the ramp angle have the same impact on the performance of the cantilevered ramp injector, and the kidney-shaped plume is formed with shorter distance with the increase of the swept and ramp angles. At the same time, the shape of the injectant mole fraction contour at X/H=6 goes through a transition from a peach-shaped plume to a kidney-shaped plume, and the cantilevered ramp injector with larger swept and ramp angles has the higher mixing efficiency and the larger drag force. The length of the step has only a slight impact on the drag force performance of the cantilevered ramp injector. However, it makes a difference to the flow field in the vicinity of the fuel injector, and the subsonic region becomes narrower with the increase of the length of the step.     

Drag reduction research in supersonic flow with opposing jet

A CFD study on drag reduction in supersonic flow with opposing jet has been conducted. Flowfield parameters, reattachment point position and surface pressure distributions are obtained and validated with experiments. From the analysis on the physical mechanism of drag reduction, it shows the phenomenon that, when the opposing jet blows, the high pressure region is located between the bow shock wave and the Mach disk, which makes the nose region much lower pressure. As the pressure ratio increases, the high pressure region is gradually pushed away from the surface. Larger the total pressure ratio is, the lower of the drag coefficient is. To study the effect of the intensity of opposing jet more reasonably, a new parameter R_PA has been introduced by combining the flux and the total pressure ratio. The study shows that the same shock wave position and drag coefficient can be obtained with the same R_PA with different fluxes and the total pressures, which means the new parameter could stand for the intensity of opposing jet and could be used to analyze the influence of opposing jet on flow field and aerodynamic force.     

Transient analysis of counterflowing jet over highly blunt cone in hypersonic flow

Understanding the characteristics of various Counterflowing jets exiting from a nose cone is crucial for determining heat load reduction and usage of this device in various conditions. Such jets can undergo several flow regimes during venting, from initial supersonic flow, to transonic, to subsonic flow regimes as the pressure of jet decreases. A bow shock wave is a characteristic flow structure during the initial stage of the jet development, and this paper focuses on the development of the bow shock wave and the jet structure behind it. The transient behavior of a sonic counterflow jet is investigated using unsteady, axisymmetric Navier–Stokes solved with SST turbulence model at free stream Mach number of 5.75. The coolant gas (Carbon Dioxide and Helium) is chosen to inject into the hypersonic air flow at the nose of the model. The gases are considered to be ideal, and the computational domain is axisymmetric. The jet structure, including the shock wave and flow separation due to an adverse pressure gradient at the nose is investigated with a focus on the differences between high diffusivity coolant jet (Helium) and low diffusivity coolant jet (CO₂) flow scenarios.     

Igniter jet dynamics in solid fuel ramjets

The dynamics of a two dimensional plane jet injected at the base of a step, parallel to the wall, in backward facing step flow geometry is numerically studied. The objective of this work is to gain insight into the dynamics of the igniter flow field in solid fuel ramjet motors. Solid fuel ramjets operate by ingestion of air and subsequent combustion with a solid fuel grain such as polyethylene. The system of governing equations is solved with a finite volume approach using a structured grid in which the AUSM⁺ scheme is used to calculate the convective fluxes. The Spalart and Allmaras turbulence model is used in these simulations. Experimental data have been used to validate the flow solver and turbulence model simulation results. The comparison of the numerical results and experimental data has validated the use of the adopted turbulence model for the study of this type of problem. A special attention is paid to the igniter jet exit location. It is shown that the wall jet igniter, issuing from the base of the step, drastically changes the structure of recirculating region of backward facing step flow and produces large and damaging shear stress on the fuel surface. Moving the igniter jet exit location to the top of the backward facing step changes the flow field favorably, by reducing the fuel surface shear stress by an order of magnitude and maintaining the recirculating region behind the step, which can provide proper residence time for the fuel–air mixture chemical reactions.     

单喷管火箭自由喷流噪声数值模拟

利用非定常燃气流场数值模拟方法，研究超声速喷流核心区及其外围亚声速区燃气流动态分布和流动特性，在此基础上利用欧拉方程条件的伽辽金有限元方法以及FWH方法，实现并完成喷流噪声传播特性、辐射特性数值模拟。喷流噪声数值模拟结果显示：燃气流推进初期，强喷流噪声区域紧随燃气流前锋；燃气流场相对稳定后，强喷流噪声区域主要位于燃气流等能区末稍，一些小尺度试验的燃气流激波系附近也存在较强喷流噪声。这些强喷流噪声主要由燃气流前锋带动的大涡动态卷吸、燃气流强湍流脉动以及激波扰动引起。受数值模拟网格分辨率影响，当前仅能保持中低频段声压级数值模拟结果与实测结果总体接近。     

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。     

凹腔超声速燃烧室氢气燃烧流场数值模拟

对带长深比为10的凹腔结构的燃烧室二维氢燃烧流场进行数值模拟,燃料喷注方式采用凹腔上游喷注加辅加凹腔前壁、底壁、后壁喷注。采用三阶MUSCL格式求解二维含组分守恒N-S方程组,湍流模型采用剪切修正的RNGk-ε湍流模型,对喷氢燃烧工况进行了计算研究,并分别分析了凹腔中不同燃料喷注方式对燃烧特性的影响。结果表明：凹腔是火焰驻留的主要区域;凹腔上游喷注氢,可以使燃料在凹腔中混合燃烧,辅加凹腔中喷氢的三种方式对燃烧状况产生一定的影响。在凹腔前壁、底面辅加喷氢,没有增强凹腔的稳焰特性,对整个燃烧状态影响不大;在凹腔后壁喷氢,能够增加凹腔中的燃料含量,加强了回流效果,对燃烧状态影响较大。三种喷注方式都没有从根本上改变凹腔燃烧流场的特性。     

Menter转捩模型在高超声速流动模拟中的改进及验证

通过压力梯度参数和湍流普朗特数的修正,对Menter等构造的低速经验关系式转捩模型进行了改进,将基于局部信息的Menter转捩模型用于高超声速流动转捩数值模拟。在多个高超声速尖锥转捩流动模拟中,对改进后模型的可靠性进行了检验,算例考核了不同壁面温度、来流湍流度、流向压力梯度等多种流动条件。结果表明,数值计算和风洞试验的壁面温度、温度恢复因子和热流等符合得较好,修正后模型较好地模拟了转捩起始位置和转捩区长度,在高超声速边界层转捩预测中具有一定潜力。     

液体亚燃燃烧室点火装置工作特性数值研究

基于气体动力学和计算流体力学的相关理论,采用CFD-ACE＋流场计算软件,对液体亚燃燃烧室点火装置单独工作时的稳态流场进行了数值模拟.在试验验证的基础上分析了点火器室压、点火导管内径和导管的结构形式对火焰点火性能的影响.结果表明：当点火器室温和燃气流量恒定时,若保持管道的扩张比不变,选用较低室压的点火器更利于点火;在一定范围内增大导管内径可以提高火焰的点火性能;燃气在直管内的流动损失较小,出口射流的速度较高,穿透深度较大,带弯头的点火导管出口火焰特征类似,有无弯头对火焰的影响很大,而角度差异产生的影响很小.     

超声速内流场横向喷流的流动与混合特性实验研究

为研究超声速内流场中横向喷流的流动与混合特性,将丙酮蒸汽加入喷流介质,用平面激光诱导荧光(PLIF)技术对流场中流向中心截面和横截面上的丙酮进行成像,研究了喷流的运动轨迹、流场结构、混合方式,以及参数对喷流流动与混合的影响。结果表明:喷流柱的波动失稳及喷流剪切层中生成的大尺度结构有助于增强喷流与主流在近场的混合;提高出口马赫数会导致剪切层失稳以及出现大尺度结构的位置移向下游,不利于改善近场的混合;增大喷口直径能增加喷流在展向的扩展,升高喷流总压能增加喷流在展向和横向的扩展,并使出现大尺度结构的位置靠近上游;在喷注流量相同条件下,采用小喷注面积高总压喷注更利于增强混合。     

超声速气流中液体横向射流的气液相互作用过程数值研究

采用Euler-Lagrange方法对来流马赫数为1.94的超声速气流中液体横向射流的气液相互作用过程进行数值研究。计算给出的射流穿透深度、液滴Sauter平均直径(SMD)及液滴速度分布均与实验吻合较好。仿真结果较详细地揭示液体射流喷雾与气流之间的强烈相互作用过程。受液雾影响,射流前形成较强激波,气流依次经过激波及液雾区域,气流速度存在两次下降过程。计算结果揭示,超声速来流可以与射流的液滴轨迹相交,气流经液雾前沿进入液雾区域后,流向往壁面偏折。本文首次发现并提出,由于气液相互作用诱导形成两组反向反转漩涡对,这对于理解两相混合过程具有重要意义。气液相对滑移速度的分析表明,液滴在穿透自由来流并开始转向时受气流作用最为显著,完成转向后气液相互作用逐渐减弱。     

燃烧室缝槽气膜冷却方案研究

针对超音速飞行器冲压发动机高马赫数、长航时的特点,结合工程计算方法和设计思想,建立了燃烧室缝槽气膜冷却过程一维计算模型,详细研究了各主要因素对气膜冷却效果的影响,并给出了某型冲压发动机高温燃烧室缝槽气膜冷却结构参考设计方案。结果表明,通过改善结构布局,合理分配缝隙冷气流量,可以有效地提高气膜冷却效果、降低壁温,适应高温燃气参数分布对隔热屏的热防护要求。     

激波在狭缝中绕射过程研究

针对固体火箭发动机药柱存在裂纹的情况，对运动激波在狭缝中绕射传播的非定常流场进行了计算，得到了清晰的流场变化图谱，从流场图中可分辨出激波在狭缝中的传播过程。对不同深宽比的狭缝进行研究，发现流场结构存在明显差别，在深宽比较小时狭缝内压力波振荡较小，当深宽比超过一定值时狭缝内的压力波产生剧烈振荡。研究结果表明，该方法可很好地描述激波在狭缝中的传播过程。