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.

A New Simulation Method for 3D Propellant Grain Burn Analysis of Solid Rocket Motor

Describes a new computer program (Regress-3D) to simulate the regression of complex 3D grain cavity and calculate the burning surface area. It has a large region of applicability in solid rocket motor design and has made new improvements compared with other available codes. User can easily and rapidly build his initial grain shapes and then obtain geometric information of his design. Considering with the calclulting results, redesigning can be performed as desire until reaching at the satisfied result. Advantages and disadvantages of this method are also discussed.

Overview of Al-based nanoenergetic ingredients for solid rocket propulsion

The introduction of nano-sized energetic ingredients first occurred in Russia about 60 years ago and arose great expectations in the rocket propulsion community, thanks to the higher energy densities and faster energy release rates exhibited with respect to conventional ingredients. But, despite intense worldwide research programs, still today mostly laboratory level applications are reported and often for scientific purposes only. A number of practical reasons prevent the applications at industrial level: inert native coating of the energetic particles, nonuniform dispersion, aging, excessive viscosity of the slurry propellant, possible limitations in mechanical properties, more demanding safety issues, cost, and so on.This paper describes the main features in terms of performance of solid rocket propellants loaded with nanometals and intends to emphasize the unique properties or operating conditions made possible by the addition of the nano-sized energetic ingredients. Steady and unsteady combustion regimes are examined.

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.

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.

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.

Study on the Effect of Solid Particles on Jet Flow and Base Thermal Environment for Solid-Liquid Bundled Rocket

Compared with using liquid rocket engines,there are a lot of high-temperature solid particles in the solid-liquid bundled rocket,which make the rocket base thermal environment worse.In order to study the influence of high-temperature solid particles on the base thermal environment,firstly,the effect of particle diameter on the jet distribution and the thermal environment in a single solid motor jet was analyzed using a numerical simulation method,and the results were compared with those of a ground test.Further,the effects of high-temperature solid particles on the jet and the thermal environment of the solid-liquid bundled rocket were analyzed and compared with flight data.The results show that high-temperature solid particles can increase the jet temperature and reduce the jet velocity.The larger the particle diameter,the greater the impact on the jet core temperature.The role of high-temperature solid particles cannot be ignored in the study of the base thermal environment.

Space Elastic Modal Modeling and Control of A New Generation Solid Bundled Rocket

In this paper,based upon the characteristics of elastic modal combination of large solid bundled launch vehicles,the finite element theory is used to describe the complex elastic vibration of a solid bundled launch vehicle,and a new three-channel unified elastic vibration equation was established.The elastic vibration equation can reflect the coupling between channels and between boosters and core stage.Some issues need consideration in the theoretical derivation,an engineering application was proposed,and the elastic vibration model was verified.The results of the theoretical derivation and simulation show that the elastic vibration equation of a solid bundled launch vehicle established in the paper is correct and can meet the needs for the engineering application.

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

长时间工作固体火箭发动机长尾喷管外绝热装置设计

固体火箭发动机长尾喷管周围一般安装有电子仪器,故而对喷管壳体表面温度有严格的限制。对于某长时间工作的发动机,为满足喷管壳体表面温度要求,同时降低消极质量,设计了一种新型双腔外绝热装置,以实现在较小的消极质量下具有较好的绝热性能。对该绝热装置进行了内流场及传热计算分析,而后进行了绝热结构优化及试验验证。结果表明:该双腔外绝热装置可以将长时间工作固体火箭发动机长尾喷管外壁温度从483℃降低至103℃,降低约78.7%;而采用同等材料同等质量的常规外防护装置只能从483℃降低至227℃,降低约53.0%;而要达到相同的防护效果,同等材料常规外防护装置的质量将增加约160%。双腔外绝热装置中段及尾端处温度随时间变化趋势与实测结果吻合较好,有望应用于其他长时间工作发动机。

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。