Improvement in Time Efficiency in Numerical Simulation for Solid Propellant Rocket Motors(SPRM)

The main purpose of the present work is to study the possibilities of reducing calculation time while maintaining the validity in the numerical simulation of the combustion product flow in SPRM chamber.Three ways of decreasing the calculation time-the use of numerical methods of high accuracy order,the reduction in spatial dimension of the problem,and the use of physical features of the processes in SPRM chamber while constructing a calculation model-were considered.Presented calculation data show that the use of these approaches makes it possible to reduce the time for solving the problems of SPRM simulation significantly(up to 100times).Also conclusions about the applicability of the mentioned above approaches in SPRM design were made.

The properties of Sn-Zn-Al-La fusible alloy for mitigation devices of solid propellant rocket motors

The Al and La elements are added to the Sn9Zn alloy to obtain the fusible alloy for the mitigation devices of solid propellant rocket motors. Differential scanning calorimetry(DSC), metallographic analysis,scanning electron microscopy(SEM), energy dispersive spectroscopy(EDS), tensile testing and fracture analysis were used to study the effect of Al and La elements on the microstructure, melting characteristics, and mechanical properties of the Sn9Zn alloy. Whether the fusible diaphragm can effectively relieve pressure was investigated by the hydrostatic pressure at high-temperature test. Experimental results show that the melting point of the Sn9Zn-0.8Al0·2La and Sn9Zn-3Al0·2La fusible alloys can meet the predetermined working temperature of ventilation. The mechanical properties of those are more than 35% higher than that of the Sn9Zn alloy at-50°C-70°C, and the mechanical strength is reduced by 80% at 175°C. It is proven by the hydrostatic pressure at high-temperature test that the fusible diaphragm can relieve pressure effectively and can be used for the design of the mitigation devices of solid propellant rocket motors.

China's Civil Solid Rocket Motors

China develops solid fuelrocket technology completely de-pending on its own strength.Since the late 1950s,variouskinds of solid rocket engines de-veloped by the 4th Academy ofChina Aerospace Corporation(CASC)have been used intomany positioning and recoveringtests of various satellites and over100 flight tests of various missileswith a 100％ successful rate.InOctober,1982,especially,Chinasuccessfully launched a solid fuelrocket from a submarine underwater,which reflected China’shigh level in this field.China’scivilian solid rocket engines are asfollows:

Solid Rocket Motors for Launch Escape Tower Provide Safety to Astronauts

Compared with other ordinary launch vehicles, the LM-2F launch vehicle has two special systems, the fault detection processing system and launch escape system, because it is used for manned spaceflight missions. During the 10 flight missions carried out by LM-2F, only the

Study on Instable Combustion of Solid Rocket Motor with Finocyl Grain

The instable combustion or oscillation combustion which occurs in three high capacity solid rocket motors using high energy composite propellant with finocyl grain is studied. The reasons of the acoustic combustion instability are also discussed. Three engineering methods that can eliminate combustion instability are proposed and discussed. The study shows that the combustion instability mainly depends on the propellant grain shape and nozzle structure. Some measures to reduce the acoustic energy and mass generation rate of combustion gas can be adopted. The test results indicate that the modified rocket motors can significantly eliminate the instable combustion and improve the motor internal ballistic performance.

Measurement of alumina film induced ablation of internal insulator in solid rocket environment

This study investigates the ablation of internal insulation induced by the deposition of an alumina film with different lateral film speeds.A sub-scale test solid rocket motor(SRM)was designed in an impinging jet configuration to form an alumina film on the sample and to encourage the lateral movement of the film by a high-speed wall jet.Fifteen static fire tests of the test SRM were conducted with six different jet velocities(V_(jet)=100 m/s,150 m/s,200 m/s,268 m/s,330 m/s,and 450 m/s)that indirectly affected the velocity of the wall jet and the deposition rate of alumina droplets.The ablation velocity was deduced from the difference in the sample thickness after a test using a coordinate measuring machine.The droplet deposition mass flux and wall jet velocity were obtained via two-phase flow simulation with the same jet velocity and effective pressure.As a result,the characteristics of alumina-induced ablation and the changes in ablation with jet velocities were obtained.The area within0.8×jet diameter was focused upon,where the ratio of ablation velocity to incoming alumina mass was constant for each jet velocity,and showed a similarity in jet structure.When the ablation velocity was increased from 2.05 to 9.98 mm/s with increasing jet velocity,the ratio of the ablation velocity and alumina mass flux decreased from 1.07×10^(-4)to 0.49×10^(-4)m^(3)/kg as Al_(2)O_(3)-C reactions became less efficient with a reduced residence time of the film.Because the decrease in residence time by the wall jet is more pronounced for slow reactions involved in Al_(2)O_(3)-C reactions,fast reactions in Al_(2)O_(3)-C reactions are less affected and result in a convergence of the volumetric rate of ablation per unit mass of alumina.

Research on the Initial Ignition of the Underwater Launching Solid Rocket Motor

The aim of this investigation is to research the initial ignition of the underwater-launching solid rocket motor.The MIXTURE multiple-phase model was set to simulate the initial ignition.The water vaporization was researched and the energy transfer was added to the energy equations.The flow field and the vaporization were calculated coupled.The initial ignition process of the underwater solid rocket motor is obtained and the vaporization influence to the underwater launching is analyzed.The "neck","inverted jet" and "eruption" phenomenon of the bubble are observed.The bubble increases more rapidly because the steam mass added to the fuel.The temperature is lower considering the vaporization because the steam enthalpy is lower than the fuel enthalpy and the flow field of the initial ignition of the underwater-launching solid rocket motor is accordant well to the reference.

Influence of Solid Rocket Motor Slag on the Space Debris Environment

The resulting slag particles from solid rocket motor( SRM) firings are an important component of space debris environment. Slag sizes as large as 1 cm have been witnessed in ground tests,and comparable sizes have been also estimated via observations of sub-orbital tail-off events. We achieve slag initial data based on MASTER slag model and SRM historical launch data,and propagate slag long-term orbital evolution taking into account the zonal harmonics J2,atmospheric drag,solar radiation pressure and luni-solar attraction to discuss the slag size distribution and orbital characteristics. Finally,future slag debris environment is evaluated based on two different launch rate assumptions. The result shows that current launch frequency will make the slag population sustain growth and the population will not decrease at once even if there are no more launches in the future.

Laminar and Turbulent Characteristics of the Acoustic/Fluid Dynamics Interactions in a Slender Simulated Solid Rocket Motor Chamber

In this paper,analytical,computational,and experimental studies are integrated to examine unsteady acoustic/vorticity transport phenomena in a solid rocket motor chamber with end-wall disturbance and side-wall injection.Acoustic-fluid dynamic interactions across the chamber may generate intense unsteady vorticity with associated shear stresses.These stresses may cause scouring and,in turn,enhance the heat rate and erosional burning of solid propellant in a real rocket chamber.In this modelling,the unsteady propellant gasification is mimicked by steady-state flow disturbed by end-wall oscillations.The analytical approach is formulated using an asymptotic technique to reduce the full governing equations.The equations that arise from the analysis possess wave properties are solved in an initial-boundary value sense.The numerical study is performed by solving the parabolized Navier–Stokes equations for the DNS simulation and unsteady Reynolds-averaged Navier–Stokes equations along with the energy equation using the control volume approach based on a staggered grid system with the turbulence modelling.The v2-f turbulence model has been implemented.The results show that an unexpectedly large amplitude of unsteady vorticity is generated at the injection side-wall of the chamber and is then penetrated downstream by the bulk motion of the internal flow.These stresses may cause a scouring effect and large transient heat transfer on the combustion surface.A comparison between the analytical,computational,and experimental results is performed.

On the Transient Numerical Simulation of Solid Rocket Motor by Coupling Quasi One⁃Dimension Internal Flow with Three⁃Dimension Propellant Grain Burnback

The quasi one⁃dimension compressible flowfield coupled to the three⁃dimension propellant grain regression solved by the level⁃set method was used to simulate the transient internal ballistics of solid rocket motor.One⁃dimension flowfield instead of three⁃dimension can save computational cost on the premise of calculation accuracy because the radial and azimuthal variations parameters have little contribution to the internal flowfield.The grain regression in real⁃time could provide accurate geometrical information for simulation.A combination of flowfluid solver and grain regression can reappear in a relatively real internal ballistic flowfield,so it is good for further studying the instability of solid rocket motor.For level⁃set equations,the total variation⁃diminishing second⁃order Runge⁃Kutta method for temporal derivatives and a fifth⁃order weighted⁃essentially⁃non⁃oscillatory scheme for spatial derivatives were used.The total variation⁃diminishing MacCormack method was used to discrete the Euler equations in flowfield solver.Two modules of this code were tested in this study:one is the burning rate module and the other is the nozzle erosion module.Results show that the burning rate influenced the solid rocket motor efficiency,and the velocity profile in the chamber was affected by the nozzle shape,and the nozzle erosion could influence the head⁃end pressure spike.

Research on Instantaneous Thrust Measurement for Attitude-control Solid Rocket Motor

In order to measure the instantaneous thrust of a certain attitude-control solid rocket motor, based on the analysis of the measurement principles, the difference between the instantaneous thrust and steady thrust measurements is pointed out. According to the measurement characteristics, a dynamic digital filter compensation method is presented. Combined the identification-modeling, dynamic compensation and simulation, the system's dynamic mathematic model is established. And then, a compensation digital filter is also designed. Thus, the dynamic response of the system is improved and the instantaneous thrust measurement can be implemented. The measurement results for the rocket motor show that the digital filter compensation is effective in the instantaneous thrust measurement.

SRM用C/C复合材料抗烧蚀技术发展

C/C复合材料因其优异的高温力学性能和抗烧蚀、抗冲击性能,已被广泛用作SRM的喉衬。本文详细介绍和分析了C/C复合材料的热化学烧蚀机理和机械剥蚀机理,在此基础上综述了近年来提高C/C复合材料抗烧蚀性能的方法及其研究进展,并对各种方法的优缺点进行分析,依据各种方法的优劣并结合工程实践对国内抗烧蚀C/C喉衬工程化应用发展提出建议。

节流式燃/氧分离组合固体发动机推力性能调节数值研究

节流式燃/氧分离组合固体发动机将燃烧室分为富燃燃烧室和富氧燃烧室,采用富燃固体推进剂和富氧固体推进剂串联装填的形式,在两个燃烧室中间增加节流阀,通过调节节流阀开度实现推力调节。利用商业软件对节流式燃/氧分离组合固体发动机在节流阀流道半径不同时的工作状态进行模拟,研究节流式燃/氧分离组合固体发动机的工作特性,获得节流式燃/氧分离组合固体发动机燃气流动和掺混特征,揭示节流式燃/氧分离组合固体发动机推力调节机理以及节流阀流道半径对节流式燃/氧分离组合固体发动机燃烧效率的影响规律。研究发现:节流阀流道开度(即节流阀流道半径R)减小可增大发动机推力;随着节流阀开度的减小,节流阀出口燃气马赫数随之增大,而发动机燃烧效率却出现先增大后减小的规律,当节流阀出口燃气流动达到临界状态时发动机燃烧效率出现最大值。

节流式燃/氧分离组合固体火箭发动机动态工作过程数值研究

利用商业软件ANSYS Fluent对节流式燃/氧分离组合固体火箭发动机动态工作过程进行数值研究,揭示该发动机基本动态工作过程以及节流阀作动速度的影响。结果表明,节流阀作动会导致发动机富氧燃烧室压强、节流阀出口质量流量以及发动机推力出现负调现象;节流阀作动速度差异不影响发动机调节后的稳定工作状态;随着节流阀作动速度增大,发动机达到稳定工作状态所需时间逐渐缩短,实现性能调控的响应速度增快,但发动机各参数负调量也随之增大,从而增加了发动机发生振动的风险。

后摆心潜入式喷管柔性接头力学特性数值仿真

针对后摆心潜入式喷管柔性接头在燃气压力与摆动联合载荷下,弹性件与增强件的力学特性计算问题,建立了对称刚体-柔性体结合的几何模型;采用了有限元方法,获取了柔性接头在燃气压力、摆动载荷及联合载荷下应力、弹性比力矩分布规律。研究表明:随着燃气压力增大,弹性比力矩降低;随摆角增大,弹性比力矩增大。文中的数值仿真方法,可以应用于柔性接头力学特性的计算,为柔性接头强度的安全性评估提供指导。

强冲蚀条件下SRM绝热层烧蚀特性实验分析

利用高过载收缩管实验发动机,开展了强冲蚀条件下的绝热层烧蚀特性研究。实验结果表明:颗粒冲蚀区的炭化层厚度明显薄于非冲蚀区,且炭化层抗剪切能力较弱。通过机理分析,认为高浓度颗粒流的强冲蚀对炭化层具有强烈的剥蚀作用。结合数值计算方法,发展了一种利用少量烧蚀实验结果,获取大量烧蚀规律和信息的数据分析方法。建立了实验条件下的高硅氧酚醛材料的烧蚀量和颗粒冲蚀状态之间的函数形式,并给出了颗粒冲蚀状态与烧蚀率之间的经验公式,具有较强的工程实用价值。

SolidWorks二次开发在燃面计算中的应用

以SolidWorks二次开发接口为基础,探讨了参数化设计在固体火箭发动机装药燃面计算中的应用.通过不同的药柱建模方法实现了固体药柱的装药计算,发展了一种更为简单准确的燃面计算方法.计算结果表明,理论计算和试验结果吻合得较好,能够较好地模拟固体药柱的燃烧过程,可以满足工程计算上的需要,为固体装药的燃面计算提供了新的思路和方法.

基于植入式光纤传感网络的固体火箭发动机结构健康监测

为实现对固体火箭发动机结构健康状态的实时监测和评估,设计并研制了一种高强度飞秒光栅传感网络,重点研究了传感器与信号传输链路的涂覆和封装。开展了发动机界面脱粘模拟实验。设计可用于发动机应变、温度和载荷的模拟演示系统,建立了基于光纤传感器采集数据的发动机数字孪生模型,形成发动机实物与数字孪生体的精准映射。最后将光纤传感器阵列植入发动机,对发动机开展水压监测实验、高低温存储实验、长周期监测实验和点火实验。对监测结果进行无损检测验证,结果表明该传感网络能够准确测量出发动机的应变状态,为固体火箭发动机的结构健康管理提供有效的数据支撑。

喷管两相湍流数值模拟及湍流模型性能评估

研究固体火箭发动机喷管中颗粒质量流率对两相流的影响规律以及不同RANS模型对该问题的预测性能,可为喷管设计等工程应用提供重要参考。在欧拉-拉格朗日(Euler-Lagrange)框架下,通过3D大涡模拟(3D LES)研究了颗粒质量流率对喷管两相流场的影响规律,以3D LES结果作为参考依据,分析了不同RANS模型针对喷管两相流在2D轴对称和3D模拟中的性能。结果表明:由于扩张段内颗粒集中在中轴线附近区域,颗粒质量流率越大,该区域马赫数越低,温度越高;壁面附近存在无颗粒区,故边界层内流动几乎不受颗粒影响。另外,采用不同湍流模型计算得到的颗粒分布与LES基本相同。在2D轴对称RANS模拟中,发现RNG k-ε模型所预测的气相质量流率和喷管比冲与LES结果吻合最好;在高颗粒质量分数下(约30%),不同湍流模型预测的相对比冲损失差别可超过3%。在3D RANS模拟中,发现Realizable k-ε模型表现出最佳的综合性能,所预测的物理量沿中轴线分布的准确性均优于其他模型。当颗粒质量分数约为31.2%时,Realizable k-ε模型所预测的比冲为与LES结果相比误差仅1.56%。