三种湍流对Scramjet燃烧室冷态流场的数值模拟

为验证从超音速混合层中发展而来的修正可压缩性(包括膨胀可压缩性和结构可压缩性两部分)的k-ε湍流模型对复杂超音速流场的预测能力,对复杂的超音速燃烧冲压发动机(Scramjet)燃烧室冷态流场进行数值模拟.将修正k-ε模型、标准k-ε模型以及标准k-ωSST模型的计算结果与实验结果进行了对比.结果表明,修正k-ε模型能够较好地预测该复杂流场,对湍流动能的预测改进较大,优于k-ωSST的结果,与实验吻合得较好;壁面压力分布以及速度分布在激波附近修正明显,与实验结果更加吻合.

MHD-Arc-Scramjet联合循环发动机的性能分析

以MHD-Arc-Scramjet联合循环发动机为对象开展性能分析，探讨了能量旁路系统对发动机性能的影响，给出了MHD-Arc-Scramjet联合循环发动机相对于传统冲压发动机的单位推力差异。结果表明：电子束电离消耗的能量，能量旁路比例，电弧注入过程的最高温度以及进气道的压缩程度是能量旁路系统中影响发动机单位推力的主要因素；在这些参数的取值范围内合理取值，并兼顾材料和技术实现等约束，性能分析表明：在来流速度2800-4000m／s范围内MHD-Arc-Scramjet联合循环发动机相对传统冲压发动机具有明显的单位推力优势。

Scramjet燃烧室流场的三维并行数值模拟及试验比较

作者采用时间相关法,通过有限体积离散,运用带化学反应的全N S方程,在神州巨型机上,针对试验模型,对油气比Φ=0.0和0.35的喷氢Scramjet燃烧室流场进行了三维并行数值模拟,得到了流场的精细结构。并行模拟所得壁面压力分布与试验所测得的壁面压力分布吻合较好。

SCRAMJET INLET MULTI-OBJECTIVE OPTIMIZATION BASED ON RESPONSE SURFACE METHODOLOGY

The uniform design and response surface methodology （RSM） are applied to the multi-objective optimization of a 2-D mixed compression scramjet inlet. The set of experimental design points on the design space is selected by the uniform design, and the inlet performance is analyzed by computational fluid dynamics （CFD）. Then complete quadratic polynomial response surface approximation models are constructed based on the performance analysis results and then used to replace theoriginal complex inlet performance model. The optimization is conducted using a multi-objective genetic algorithm NSGA-Ⅱ, and the Pareto optimal solution set is obtained. Results show that the uniform design and RSM can reduce the computational complexity of numerical simulation and improve the optimization efficiency.

Scramjet尾喷管几何调节方案的计算与实验研究

高超声速飞行器飞行接力点和巡航结束点尾喷管冷、热态俯仰力矩差较大,给飞行器的飞行姿态控制造成严重影响。为了减小喷管冷、热态俯仰力矩差,提出了在喷管上膨胀面末端增加移动板进行调节的方案,并进行了详细的三维数值模拟和相应的风洞缩比冷流实验研究。计算结果表明,Ma=4.5时,调节移动板伸出400mm,喷管冷、热态力矩差最大减小21.74%,推力系数损失1.64%;Ma=6.5时,调节移动板喷管冷、热态力矩差可降低77.59%,而推力系数只减小1.35%,调节收益非常明显。最后通过将喷管各调节状态下的冷流缩比实验壁面压力数据与计算结果的对比,证明了该调节方案的计算方法及其结果是可靠的,同时得出该调节方案可以有效地降低冷、热态力矩差的结论。

Scramjet整机流场的非结构网格并行数值模拟

为了得到某支板喷氢超燃冲压发动机内流场的细节,采用非结构网格CFD隐式解算器,耦合求解Na-vire-Stokes方程、组分连续方程和M enter的k-ω应力剪切传输(SST)湍流模型,对整机模型的冷流和三个当量油气比的湍流反应流进行了三维并行模拟。给出了流线和水的分布云图,分析了支板附近旋涡和推进流道一维质量加权平均的马赫数、静温、组分质量分数及燃烧效率等参数随当量油气比的变化,并将壁面压力和实验值进行了比较,吻合较好。除=0.8在支板附近出现了小范围亚声速外,其余状态整个内流道超声速占主流。

Experimental Study on Effects of Fuel Injection on Scramjet Combustor Performance

In order to investigate the effects of fuel injection distribution on the scrarnjet combustor performance, there are conducted three sets of test on a hydrocarbon fueled direct-connect scramjet test facility. The results of Test A, whose fuel injection is carried out with injectors located on the top-wall and the bottom-wall, show that the fuel injection with an appropriate close-front and centralized distribution would be of much help to optimize combustor performances. The results of Test B, whose fuel injection is performed at the optimal injection locations found in Test A, with a given equivalence ratio and different injection proportions for each injector, show that this injection mode is of little benefit to improve combustor performances. The results of Test C with a circumferential fuel injection distribution displaies the possibility of ameliorating combustor performance. By analyzing the effects of injection location parameters on combustor performances on the base of the data of Test C, it is clear that the injector location has strong coupled influences on combustor performances. In addition, an irmer-force synthesis specific impulse is used to reduce the errors caused by the disturbance of fuel supply and working state of air heater while assessing combustor performances.

模拟Scramjet隔离段实际入口的直连试验设备设计试验验证

为了模拟进气道喉道非均匀流场对超燃冲压发动机隔离段流动特性的影响,针对典型超燃冲压发动机隔离段入口流场的马赫数分布要求,设计了隔离段直连试验设备。该设备整个通道采用扩张构型,在附面层发展段的下壁面采用加圆形凸台的方法加速附面层发展,并针对该设备进行了数值模拟和冷流试验。结果表明,该直连试验设备不仅能准确模拟隔离段实际入口流场(即入射激波和非对称附面层),模拟的主流区马赫数与目标值差值在0.02以内,还解决了传统的直连试验设备存在的启动问题。试验与数值模拟结果吻合较好,验证了该直连试验设备设计方法的可行性。

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.

Simulation of turbulent combustion in DLR Scramjet

Turbulent combustion in a DLR (German Aerospace Center) Scramjet engine was simulated using the newly-proposed Partially Resolved Numerical Simulation (PRNS) procedure. The PRNS procedure uses temporal filtering to define large-scale turbulence, and the model developed to account for unresolved scales is grid independent. No problem about inner commutation error and inconsistencies will arise from the PRNS, while such issues are of concern in traditional Large Eddy Simulation (LES) methods. The mean results have good agreement with the experiment data and the flow structures with small scales are well resolved.

Solution Method for Singular Initial Value Problems of One-dimensional Steady Transonic Flow in a Dual-mode Scramjet

Singular initial value problems arise in solving one-dimensional steady transonic flow of dualmode scramjet. The existing solution method has the problems of large initial value errors in principles. This paper puts forward an improved algorithm based on variable transformation, and constructs a nonsingular one-dimensional steady transonic flow equation by defining a new variable. The improved algorithm can eliminate the singularity of the differential equation, and can solve the singular initial value problems of one-dimensional steady transonic flow of dual-mode scramjet.

Effects of Wall Emissivity on Aerodynamic Heating in Scramjets

The effects of the wall emissivity on aerodynamic heating in a scramjet are analyzed.The supersonic turbulent combustion flow including radiation is solved in the framework of a decoupled strategy where the flow field is determined first and the radiation field next.In particular,a finite difference method is used for solving the flow while a DOM(iscrete ordinates method)approach combined with a WSGGM(weighted sum of gray gases)model is implemented for radiative transfer.Supersonic nonreactive turbulent channel flows are examined for a DLR hydrogen fueled scramjet changing parametrically the wall emissivity.The results indicate that the wall radiative heating rises greatly with increasing the wall emissivity.As the wall emissivity rises,the radiative source and total absorption increase,while the incident radiation decreases apparently.Notably,although the radiative heating can reach a significant level,its contribution to the total aerodynamic heating is relatively limited.

Fabrication and Thermal Structural Characteristics of Ultra-high Temperature Ceramic Struts in Scramjets

ZrB_2-SiC based ultra-high temperature ceramic（UHTC） struts were firstly proposed and fabricated with the potential application in the combustor of scramjets for fuel injection and flame-holding for their machinability and excellent oxidation/ablation resistance in the extreme harsh environment. The struts were machined with electrospark wire-electrode cutting techniques to form UHTC into the desired shape, and with laser drilling to drill tiny holes providing the channels for fuel injection. The integrated thermal-structural characteristic of the struts was evaluated in high-temperature combustion environment by the propane-oxygen free jet facility, subject to the heat flux of 1.5 MW/m^2 lasting for 300 seconds, and the struts maintained integrity during and after the first experiment. The experiments were repeated for verifying the reusability of the struts. Fracture occurred during the second repeated experiment with the crack propagating through the hole. Finite element analysis（FEA） was carried out to study the thermal stress distribution in the UHTC strut. The simulation results show a high thermal stress concentration occurs at the hole which is the crack initiation position. The phenomenon is in good agreement with the experimental results. The study shows that the thermal stress concentration is a practical key issue in the applications of the reusable UHTC strut for fuel injection structure in scramjets.

Engine performance analysis and optimization of a dual-mode scramjet with varied inlet conditions

A dual-mode scramjet can operate in a wide range of flight conditions. Higher thrust can be generated by adopting suitable combustion modes. Based on the net thrust, an analysis and preliminary optimal design of a kerosene-fueled parameterized dual-mode scramjet at a cru- cial flight Mach number of 6 were investigated by using a modified quasi-one-dimensional method and simulated annealing strategy. Engine structure and heat release distrib- utions, affecting the engine thrust, were chosen as analytical parameters for varied inlet conditions （isolator entrance Mach number： 1.5-3.5）. Results show that different opti- mal heat release distributions and structural conditions can be obtained at five different inlet conditions. The highest net thrust of the parameterized dual-mode engine can be achieved by a subsonic combustion mode at an isolator entrance Mach number of 2.5. Additionally, the effects of heat release and scramjet structure on net thrust have been discussed. The present results and the developed analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets.

Study on Feasibility of Reusable Rocket Launching Technology by Use of Scramjet and Maglev Technologies

The scramjet and maglev engineering technology development and trends at home and abroad are firstly presented in this paper. A new launch mode of space transportation system is proposed based on scramjet and magnetic suspension technologies, and its key technologies required are given. This paper also makes analysis on using scramjet and magnetic suspension technologies to launch a reusable rocket, and the results show that a normal temperature conductor maglev launch system is feasible.

Flame quenching process in cavity based on model scramjet combustor

The flame quenching process in combustors was observed by high speed camera and Schlieren system, at the inflow conditions of Ma = 2.64, To = 1483K, P0 = 1.65 MPa, T = 724 K and P -- 76.3 kPa. Changing process of the flame and shock structure in the combustor was clearly observed. The results revealed that the precom- bustion shock disappeared accompanied with the process in which the flame was blown out and withdrawed from the mainflow into the cavity and vanished after a short while. The time of quenching process was extended by the cavity flame holder, and the ability of flame holding was enhanced by arranging more cavities in the downstream as well. The flame was blown from the upstream to the downstream, so the flame in the downstream of the cavity was quenched out later than that in the upstream.

Experimental Study of Ethylene Combustion in a Scramjet Combustor

In this paper the ignition characteristics of gaseous ethylene hydrocarbon fuel is investigated in the supersonic clean airstreams experimental facility with a resistance heater. The generic cavity flame holder is used to create recirculation and promote the fuel/air mixing at the lower wall of the combustor. Three different injection concepts are considered in this research ： （ 1 ） ethylene injection upstream of the cavity ; （2） ethylene and hydrogen injection upstream of the cavity simultaneously; （ 3 ）ethylene injection preceded by pilot hydrogen injection. The pilot injection showed to be a supportive tool for holding the flame of the main normal ethylene fuel injection. Therefore, using pilot hydrogen injection and cavity configuration necessitates optimizing the combustor length to ensure the complete combustion and the full liberation of the chemical energy stored in the fuel before exiting the combustor.The present study proved the possibility of igniting the ethylene and maintaining its flame in the supersonic airstreams.

Effects of the nozzle structure and fluidized gas composition on the gas-particle two-phase jet characteristic in a powder fuel scramjet

The interaction between nozzle design and fluidization gas composition significantly influences the dynamics within a powder fuel scramjet's combustion chamber.To investigate this relationship,an experimental study utilized high-speed shadow imaging technology to explore the macroscopic aspects of powder fuel injection.The investigation examined various convergence angles,nozzle throat lengths,and fluidized gas compositions.Key findings include:During jet development,powder fuel initially concentrates near the axis,with non-convergence angle nozzles exhibiting longer concentrated distribution periods than convergence angle conditions.Decreasing nozzle convergence angles lead to increased penetration distance,frontal velocity,and radial diffusion distance during the initial stages of jet development.Additionally,stable jet shapes show larger divergence angles as nozzle convergence angle decreases,with the largest divergence angle observed atα=60°.In the initial 0-7 ms of jet development,the powder fuel jet demonstrates greater penetration distance and frontal velocity under certain conditions.Moreover,penetration distance and frontal velocity increase with throat length from 7 to 20 ms,accompanied by changes in divergence angles.Specifically,at a throat length(l)of 2 mm,the near-field divergence angle measures 46.50°,and the far-field divergence angle is 22.25°.Conversely,at l=8mm,the near-field divergence angle is 33.49°,and the far-field divergence angle is 23.21°.The fluidization gas composition minimally affects jet penetration distance and frontal velocity during the initial 0-3 ms.However,due to hydrogen's low density,hydrogen/powder fuel jets exhibit shorter distances and velocities compared to nitrogen/powder fuel jets.Hydrogen fluidization also results in larger divergence angles,particularly in the near field.These findings underscore the importance of nozzle design and fluidization gas composition in optimizing scramjet performance and efficiency.

Drag reduction characteristics of recirculation flow at rocket base in an RBCC engine under ramjet/scramjet mode

To reduce the drag generated by the recirculation flow at the rocket base in a RocketBased Combined Cycle(RBCC)engine operating in the ramjet/scramjet mode,a novel annular rocket RBCC engine based on a central plug cone was proposed.The performance loss mechanism caused by the recirculation flow at the rocket base and the influence of the plug cone configuration on the thrust performance were studied.Results indicated that the recirculation flow at the rocket base extended through the entire combustor,which creates an extensive range of the"low-kineticenergy zone"at the center and leads to an engine thrust loss.The plug cone serving as a surface structure had a restrictive effect on the internal flow of the engine,making it smoothly transit at the position of the large separation zone.The model RBCC engine could achieve a maximum thrust augmentation of 37.6%with a long plug cone that was twice diameter of the inner isolator.However,a shorter plug cone that was half diameter of the inner isolator proved less effective at reducing the recirculation flow for a supersonic flow and induced an undesirable flow fraction that diminished the thrust performance.Furthermore,the effectiveness of the plug cone increased with the flight Mach number,indicating that it could further broaden the operating speed range of the scramjet mode.

Numerical simulation on magnetohydrodynamic power generation channel of scramjet

The reverse magnetohydrodynamic(MHD)energy bypass technology is a promising energy redis⁃tribution technology in the scramjet system,in augmented with a power generation equipment to supply the neces⁃sary long-distance flight airframe power.In this paper,a computational model of the scramjet magnetohydrody⁃namic channel is developed and verified by using the commercial software Fluent.It is found that when the mag⁃netic induction intensity is 1,2,3,4 T,the power generation efficiency is 22.5%,22.3%,22.0%,21.5%,and decreases with the increase of the magnetic induction intensity,and the enthalpy extraction rate is 0.026%,0.1%,0.21%,0.34%,and increases with the increase of the magnetic induction intensity.The deceleration ef⁃fect of electromagnetic action on the airflow in the power channel increases with the increase of magnetic induc⁃tion intensity.The stronger the magnetic field intensity,the more obvious the decreasing effect of fluid Mach num⁃ber in the channel.The power generation efficiency decreases as the magnetic induction intensity increases and the enthalpy extraction rate is reversed.As the local currents gathering at inlet and outlet of the power generation area,total temperature and enthalpy along the flow direction do not vary linearly,and there are maximum and minimum values at inlet and outlet.Increasing the number of electrodes can effectively regulate the percentage of Joule heat dissipation,which can improve the power generation efficiency.