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.     

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.     

凹腔结构对圆形超燃冲压发动机燃烧室阻力特性影响

凹腔作为促进燃烧室中燃料与来流混合和稳定燃烧的有效手段之一,其研究已引起人们的广泛关注.采用数值模拟方法,探索了圆形超燃冲压发动机燃烧室阻力特性随凹腔结构参数的变化趋势,同时初步考察了飞行攻角对凹腔阻力特性的影响.研究发现,凹腔摩阻相比压阻很小,凹腔时燃烧室的阻力特性主要体现在其压阻上;随着后掠角的增大,热试和冷流状态...     

双模态冲压发动机燃烧室流场数值模拟研究

在三维、粘性、湍流及有化学反应的Navier-Stokes方程基础上，通过有限化学反应速率／涡扩散模型模化湍流燃烧，对以H2为燃料的双模态冲压发动机燃烧室流场进行了研究，分析了空燃比、燃料入射角、飞行马赫数对燃烧室工作模态的影响，并分析了燃烧室隔离段的作用。     

Numerical investigation on the shock wave transition in a three-dimensional scramjet isolator

The scramjet isolator, which is used to prevent the hypersonic inlet from disturbances that arise from the pressure rise in the scramjet combustor due to the intense turbulent combustion, is one of the most critical components in hypersonic airbreathing propulsion systems. Any engineering error that is possible in the design and manufacturing procedure of the experimental model, and the intense heat release in the scramjet combustor, may cause the performance of the isolator to decrease, leading to its lack of capability in supporting the back pressure. The coupled implicit Reynolds Averaged Navier–Stokes (RANS) equations and the two-equation standard k?ε turbulent model have been employed to numerically simulate the flow fields in a three-dimensional scramjet isolator. The effects of the divergent angle and the back pressure on the shock wave transition and the location of the leading edge of the shock wave train have been estimated and discussed. The obtained results show that the present numerical results are in very good agreement with the available experimental shadow-pictures, and the numerical method is more suitable for capturing the shock wave train and predicting the location of the leading edge of the shock wave train in the scramjet isolator than the present two-dimensional numerical methods. This is due to the small width-to-height ratio of the isolator and the intense three-dimensional flow structures. On increasing the divergent angle of the scramjet isolator, the static pressure along the central symmetrical line of the isolator decreases sharply. This is due to the strong expansion wave generated at the entrance of the isolator, and when the divergent angle of the isolator is sufficiently large, namely 1.5°, a zone of negative pressure is formed just ahead of the leading edge of the shock wave train. At the same time, the shock wave train varies from being oblique to being normal, and then back to oblique. With an increase in the prescribed back pressure at the exit of the scramjet isolator, the leading edge of the shock wave train moves forward towards the entrance of the isolator, and when the back pressure is sufficiently large, unstart conditions in the hypersonic inlet can take place if the shock train reaches the inlet.     

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。     

燃油分配对超燃冲压发动机的性能影响仿真分析

针对超燃冲压发动机两级燃油分配对内流道流动过程、燃烧模态、发动机性能及调节特性的影响问题,建立了发动机一维流动分析模型;对马赫数6/当量比1,马赫数6/当量比0.6,马赫数4/当量比1三种工况不同的一级/二级燃烧室燃油分配比例下的流动过程进行了仿真,并获得了不同燃油分配规律下的发动机性能.通过分析表明：超燃冲压发动机的两级燃油分配比例直接影响发动机内流道内的流动参数分布、燃烧模态及发动机比冲等性能参数.对于马赫数6/当量比1工况,当一级燃烧室的燃油分配比例为30％～70％时,可在全流道内组织纯超声速燃烧,最高比冲超过800 s;对于马赫数6/当量比0.6工况,即使将所有的燃油均在一级燃烧室喷入,流道也不会壅塞,该工况下最大比冲超过800s;对于马赫数4/当量比1工况,燃烧室内组织亚声速燃烧,最大比冲为1 031.9 s;为保证亚声速燃烧扰动不传递到燃烧室入口外,一级燃油分配比例不应过高.     

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.     

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。     

固体火箭超燃冲压发动机补燃室构型的影响分析

针对不同补燃室结构参数对固体火箭超燃冲压发动机补燃室掺混燃烧性能的影响进行研究,分析各级燃烧室的长度与扩张角度对补燃室性能的影响。采用基于密度的二阶迎风格式对补燃室掺混燃烧进行模拟,湍流模型和燃烧模型分别采用SST k-ω模型和涡团耗散模型。结果表明,提高燃烧效率与降低总压损失是相互矛盾的;燃烧效率随燃烧室长度的增加而增大,随燃烧室扩张角度的增加而减小;总压恢复系数随燃烧室长度的增加而减小,随燃烧室扩张角度的增加而增大;一级燃烧室的结构参数对燃烧效率与总压恢复系数的影响最大。当补燃室的总长与出口面积一定时,以发动机的总体性能参数作为补燃室构型的优化目标,对一、二级燃烧室长度与一、三级燃烧室扩张角度进行优化。     

空气节流对煤油燃料超燃燃烧室燃烧性能影响

采用非定常数值模拟方法研究了空气节流对煤油燃料超燃冲压发动机燃烧性能的影响,并研究了节流流量和节流撤去时间对节流效果的影响。在发动机入口马赫数2.0、静温656.5K、静压0.125MPa的条件下,无空气节流时发动机下壁面稳焰失败,壁面压力较低;有空气节流时发动机下壁面燃料稳定燃烧,壁面压力较高。空气节流可以有效地提高发动机的推力性能,可以改变发动机的燃烧模态。随着节流流量和节流撤去时间的增加,燃烧越来越剧烈,壁面压力逐渐升高,可能影响进气道的起动。节流可能促使流场产生振荡现象,通过改变节流流量也可以消除振荡现象。     

固体燃料超燃冲压发动机研究概况

对固体燃料超燃冲压发动机的应用背景、潜在优势,以及国内外研究现状和进展做了详细阐述。从固体燃料超燃冲压发动机工作原理、固体燃料类型、数值模拟以及实验研究等方面出发,论述了固体燃料超燃发动机研究的进展和难点,并对固体燃料超燃冲压发动机未来研究趋势进行了展望。研究认为:固体燃料在超声速流动下的热量分布与表面火焰传播等方面还需要深入研究,需建立不同固体燃料的受热行为模型;应用大涡模拟方法分析微尺度下流场结构并耦合固体燃料传热传质过程的可行性需进一步确认;考虑飞行参数,进气道与隔离段性能的发动机整体数值模拟工作需要进一步加强。     

固体火箭冲压发动机的若干技术问题

简述了固体火箭冲压发动机类型及工作原理,总体评价了固体火箭冲压发动机发展时快时慢的原因,弹-机一体化设计、贫氧推进剂、进气道、转级机构、补燃室等设计中应注意的问题,提出了应加强的研究工作,即开展高能低沉积燃烧产物贫氧推进剂研究;完善多种燃气流量调节装置方案,提高其可靠性;进一步开展一次燃烧和二次掺混燃烧的理论和实验研究,提高燃烧效率;尽快建立自由射流等试验研究手段,开展相关的研究工作;适时开展固体燃料冲压、固体超燃冲压及膏体冲压等发动机的研究,不断拓宽应用领域。     

Influences of H2O mass fraction and chemical kinetics mechanism on the turbulent diffusion combustion of H2–O2 in supersonic flows

Hydrogen is one of the most promising fuels for the airbreathing hypersonic propulsion system, and it attracts an increasing attention of the researchers worldwide. In this study, a typical hydrogen-fueled supersonic combustor was investigated numerically, and the predicted results were compared with the available experimental data in the open literature. Two different chemical reaction mechanisms were employed to evaluate their effects on the combustion of H₂–O₂, namely the two-step and the seven-step mechanisms, and the vitiation effect was analyzed by varying the H₂O mass fraction. The obtained results show that the predicted mole fraction profiles for different components show very good agreement with the available experimental data under the supersonic mixing and combustion conditions, and the chemical reaction mechanism has only a slight impact on the overall performance of the turbulent diffusion combustion. The simple mechanism of H₂–O₂ can be employed to evaluate the performance of the combustor in order to reduce the computational cost. The H₂O flow vitiation makes a great difference to the combustion of H₂–O₂, and there is an optimal H₂O mass fraction existing to enhance the intensity of the turbulent combustion. In the range considered in this paper, its optimal value is 0.15. The initiated location of the reaction appears far away from the bottom wall with the increase of the H₂O mass fraction, and the H₂O flow vitiation quickens the transition from subsonic to supersonic mode at the exit of the combustor.     

Size quantization in high-temperature superconducting cuprates and a link to Einstein's diffusion law

Optimum doping of high-temperature superconductors (HTSC) defines a superconducting unit volume for each HTSC. For a single-mode HTSC, e.g., a cuprate with one CuO₂ plane, the volume is given by V_sc=cx², where c is the unit cell height and x the doping distance. The experimental resistivity at T_c is connected to the structure by ρ(exp)≈c×h/(2e²). Combining this result with the classical definition of resistivity leads to an equation similar to Einstein's diffusion law x²/(2τ)=h/(2M_eff)=D, where τ is the relaxation time, M_eff=2m_e and D the diffusion constant. It has also been shown that the mean free path d=x. The Einstein–Smoluchowski diffusion relation D=μk_BT_c provides a connection to T_c.     

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。     

SMC模式下RBCC发动机4Ma工况性能仿真

为研究SMC模式下火箭混合比对RBCC发动机性能的影响规律,完成了氢/氧火箭推力室中心布局、二元定几何结构模型发动机飞行马赫数Ma0=4、高度H=17 km弹道点流场仿真,获得了不同火箭混合比(MR=2、3、4、5、6、8)及燃烧室长度的推力、比冲性能.研究表明:在火箭燃气富燃条件下(MR<8),产生了正的火箭推力增益...     

Radial profile of pressure in a storm ring current as a function of D st

Using satellite data obtained near the equatorial plane during 12 magnetic storms with amplitudes from ?61 down to ?422 nT, the dependences of maximum in L-profile of pressure (L _m) of the ring current (RC) on the current value of D _st are constructed, and their analytical approximations are derived. It is established that function L _m(D _st ) is steeper on the phase of recovery than during the storm’s main phase. The form of the outer edge of experimental radial profiles of RC pressure is studied, and it is demonstrated to correspond to exponential growth of the total energy of RC particles on a given L shell with decreasing L. It is shown that during the storms’ main phase the ratio of plasma and magnetic field pressures at the RC maximum does not practically depend on the storm strength and L _m value. This fact reflects resistance of the Earth’s magnetic field to RC expansion, and testifies that during storms the possibilities of injection to small L are limited for RC particles. During the storms’ recovery phase this ratio quickly increases with increasing L _m, which reflects an increased fraction of plasma in the total pressure balance. It is demonstrated that function L _m(D _st ) is derived for the main phase of storms from the equations of drift motion of RC ions in electrical and magnetic fields, reflecting the dipole character of magnetic field and scale invariance of the pattern of particle convection near the RC maximum. For the recovery phase it is obtained from the Dessler-Parker-Sckopke relationship. The obtained regularities allow one to judge about the radial profile of RC pressure from ground-based magnetic measurements (data on the D _st variation).     

Analyse du fonctionnement d'un propulseur a éjection modulée

ONERA developed, for studying the response of a propellant to a pressure or velocity fluctuation, an experimental rocket engine whose nozzle throat area can be modulated by a toothed disk.The paper presents a linearized theory of the functioning of this engine in the low frequency domain, i.e. when there is no wave propagation within the combuster.To describe the functioning of this motor, the Ryazantsev-Novozhilov method, which assumes that the gas response is instantaneous, is used. This analysis takes into account the erosion and radiation effects, the combustion efficiency and the thermal losses through the walls.Two particular cases are described, for two values of the Damköhler parameter $\text{D}{}_1\text{=}\text{t}{}_{\mathrm{c}}\text{t}{}_{\mathrm{th}}$, where t_c is the residence time in the combuster and t_th the characteristic thermal time of the heat penetration into the solid propellant. These two cases correspond, one to a classical propellant D₁ > 1, the other to a particular propellant of low burning rate ($\text{J}{}_{\mathrm{b}}\text{? 0.2}\text{to}\text{0.4}\text{mm s}{}^{\mathrm{?1}}$) D₁ < 1. The stability conditions are analysed as well as the pressure amplitute and phase as a function of the nozzle throat modulation frequency.Still in linearized theory, the complete solutions of the problem are presented, using a method of numerical resolution.