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.

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.

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.

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.

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.

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.

Experimental and Theoretical Research Review of Hybrid Rocket Motor Techniques and Applications

A hybrid rocket motor combines components from both solid fuel and liquid fuel rocket motors. The fuel itself is a solid grain, (often paraffin or hydroxyl-terminated polybutadiene, known as HTPB) while the oxidizing agent is liquid (often hydrogen peroxide or liquid oxygen). These components are combined in the fuel chamber which doubles as the combustion chamber for the hybrid motor. This review looks at the advances in techniques that have taken place in the development of these motors since 1995. Methods of testing the thrust from rocket motors and of measuring the rocket plume spectroscopically for combustion reaction products have been developed. These assessments allow researchers to more completely understand the effects of additives and physical changes in design, in terms of regression rates and thrust developed. Hybrid rocket motors have been used or tested in many areas of rocketry, including tactical rockets and large launch vehicles. Several additives have shown significant improvements in regression rates and thrust, including Guanidinium azotetrazolate (GAT), and various Aluminum alloys. The most recent discoveries have come from research into nano-particle additives. The nano-particles have been shown to provide enhancements to many parameters of hybrid rocket function, and research into specific areas continues in the sub-field of nano-additives for fuel grains.

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.

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.

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:

Study on the Multi-Component Test Technology of Solid Rocket Motors

This paper presents a calculation method, and composition of a five-component force test device, to resolve the requirement for solid rocket motor multi-component force test. It compares the difference between two classical algorithms, and summarizes the structure and practical usage of the solid rocket motor multi-component force using soft connection structure measurement. The decoupling method of test data is described, and a reasonable research method for solid rocket motor multi-component force is also given in this paper.

Development of Solid Rocket Motor for Manned Launch Vehicle Escape System

The solid rocket motors for the escape system of China’s LM-2 F manned launch vehicle are described,the key technologies and technical innovations utilized are summarized.The technical features and development of foreign manned launch abort systems are also presented.The development trends of the solid rocket motor for future Chinese manned launch vehicle escape systems are proposed,which can provide a reference for the future development of manned launch vehicle escape systems.

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

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。

固体发动机喷管热结构耦合数值仿真研究

针对在热载荷与内压联合作用下,固体火箭发动机潜入式喷管的热结构耦合数值仿真问题,采用流体仿真软件与结构计算软件,展开了数值模拟。在流体仿真软件中,计算了潜入式喷管工作过程的内部稳态流场,提取了高温燃气流动参数与固体壁面对流换热系数。在有限元结构计算平台内,编写了非均布壁面压力载荷与非均布对流换热系数子程序,耦合求解了喷管热结构耦合问题,获取了热防护材料内部温度场与应力场分布。研究表明,热载荷是引起喉衬应力的主要原因。喉衬环向压应力与拉应力随时间呈现先增大后减小的过程。经地面喷管结构完整性热试车试验验证,该热结构耦合数值仿真方法可用于喷管结构完整性与安全性分析,为喷管结构设计提供合理指导。

基于RBF代理优化的固体火箭发动机喷管型面设计

针对固定扩张比与扩张段长度的二维轴对称固体火箭发动机喷管进行扩张段型面优化,优化目标为喷管推力最大化,优化参数为贝塞尔曲线控制点的径向位置,优化方法采用径向基函数(Radial Basis Function, RBF)代理优化算法。采用纯气相与两相流两种模型分别进行优化设计,纯气相的结果表明,对于10个控制点表达的贝塞尔曲线,优化后的推力提高了1.64%。以此优化型面为初始型面,增加控制点个数至16个,二次优化后的推力又提高了0.095%。增大优化参数范围,同时引入判断拐点的约束,对于10个控制点表达的贝塞尔曲线进行单轮优化,结果同上述经过两轮优化之后的结果相近,优化后的喷管推力提高了1.78%,说明算法具有较强的稳定性。通过对不同控制参数个数的贝塞尔曲线优化过程的对比,给出了合理选择控制点个数的方法与建议。两相流的优化结果表明,由于颗粒的滞后影响造成了两相流损失,两相流喷管的推力小于纯气相喷管,但两相流喷管优化后的推力较优化前初始型面的推力提高了1.87%,略高于纯气相喷管。RBF代理优化算法适用于由任意数量控制点组成的贝塞尔曲线表达的喷管扩张段型面优化,并有较高的效率与较强的稳定性。