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 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.

Numerical and experimental research on axial injection end-burning hybrid rocket motors with polyethylene fuel

This study investigates the end-burning hybrid rocket motors with polyethylene fuel by the numerical simulation and experiment.Based on computational fluid dynamics,a numerical model is developed.The model is validated by two firing tests in this hybrid rocket motor,which uses oxygen and polyethylene as propellants.The results show that the numerical and experimental data are in good agreement,and the error of the chamber pressure is less than 2.63%.Based on the simulation mode,the blowoff limit of the end-burning hybrid rocket motors is investigated.When the nozzle throat diameter and the inner diameter of grain are large,it is more difficult for the hybrid rocket motor to achieve end-burning mode,i.e.,the flame spreading is prevented in the narrow duct.The main reason is that when the nozzle throat and the grain port are large,chamber pressure and oxidizer flow velocity are low.Therefore,the friction velocity considering the pressure and flow velocity is proposed.The critical friction velocity is about 4.054–4.890 m/s in the hybrid rocket motors.When the friction velocity exceeds the critical friction velocity,the combustion mode in hybrid rocket motors changes from the flame spreading mode to the end-burning mode.Moreover,the regression rate formula is obtained by fitting,which shows that the regression rate has a good correlation with combustion chamber pressure.The critical friction velocity and regression rate formula can provide an important reference for end-burning hybrid rocket motors.

A comprehensive fluid-solid coupling dynamic simulation for spatiotemporal distribution of regression rate in hybrid rocket motors

The spatiotemporal distribution characteristics of the regression rate are crucial aspects of the research on Hybrid Rocket Motor(HRM).This study presents a pioneering effort in achiev-ing a comprehensive numerical simulation of fluid dynamics and heat transfer in both the fluid and solid regions throughout the entire operation of an HRM.To accomplish this,a dynamic grid tech-nique that incorporates fluid-solid coupling is utilized.To validate the precision of the numerical simulations,a firing test is conducted,with embedded thermocouple probes being used to measure the inner temperature of the fuel grain.The temperature variations in the solid fuel obtained from both experiment and simulations show good agreement.The maximum combustion temperature and average thrust obtained from the simulations are found to deviate from the experimental results by only 3.3%and 2.4%,respectively.Thus,it can be demonstrated that transient numerical simu-lations accurately capture the fluid-solid coupling characteristics and transient regression rate.The dynamic simulation results of inner flow field and solid region throughout the entire working stage reveal that the presence of vortices enhances the blending of combustion gases and improves the regression rate at both the front and rear ends of the fuel grain.In addition,oscillations of the regression rate obtained in the simulation can also be well corresponded with the corrugated surface observed in the experiment.Furthermore,the zero-dimension regression rate formula and the for-mula describing the axial location dependence of the regression rate are fitted from the simulation results,with the corresponding coefficients of determination(R°)of 0.9765 and 0.9298,respectively.This research serves as a reference for predicting the performance of HRM with gas oxygen and polyethylene,and presents a credible way for investigating the spatiotemporal distribution of the regression rate.

Design and optimization of variable thrust hybrid rocket motors for sounding rockets

Besides safety and low-cost,the start/shutdown/restarting and throttling ability are the other two significant advantages of hy-brid rocket motors(HRMs) compared with liquid and solid ones.In this study,a two-stage variable thrust and non-toxic 98%HP/HTPB hybrid rocket motor(VTHRM) is designed and applied in a sounding rocket,and the design parameters of the motor are analyzed and optimized.A computational program is developed to design the motor system structure,to predict the interior ballistics and the ballistic trajectory.A star grain and a wheel grain are compared.The design of experiment(DOE),variance analysis and the main effect analysis are employed to investigate the influence of the main design parameters on mo-tor performance.The multidiscipline feasible(MDF) approach is applied to establish the optimization procedure after analyz-ing the system design structure matrix.A modified differential evolution algorithm is employed to maximize the load mass.The results indicate that the wheel grain could obtain a larger load mass and a lower length to diameter ratio,and that throttling markedly meliorates the motor and rocket performance.The conclusions drawn from the analysis and optimization could pro-vide instructive guide and theoretical basis for engineering designs.

Numerical study on similarity of plume infrared radiation between reduced-scale solid rocket motors

This study seeks to determine the similarities in plume radiation between reduced and full-scale solid rocket models in ground test conditions through investigation of flow and radiation for a series of scale ratios ranging from 0.1 to 1. The radiative transfer equation (RTE) considering gas and particle radiation in a non-uniform plume has been adopted and solved by the finite volume method (FVM) to compute the three dimensional, spectral and directional radiation of a plume in the infrared waveband 2-6 mu m. Conditions at wavelengths 2.7 mu m and 4.3 mu m are discussed in detail, and ratios of plume radiation for reduced-scale through full-scale models are examined. This work shows that, with increasing scale ratio of a computed rocket motor, area of the high-temperature core increases as a 2 power function of the scale ratio, and the radiation intensity of the plume increases with 2-2.5 power of the scale ratio. The infrared radiation of plume gases shows a strong spectral dependency, while that of Al2O3 particles shows spectral continuity of gray media. Spectral radiation intensity of a computed solid rocket plume's high temperature core increases significantly in peak radiation spectra of plume gases CO and CO2 center dot Al2O3 particles are the major radiation component in a rocket plume. There is good similarity between contours of plume spectral radiance from different scale models of computed rockets, and there are two peak spectra of radiation intensity at wavebands 2.7-3.0 lm and 4.2-4.6 lm. Directed radiation intensity of the entire plume volume will rise with increasing elevation angle. (C) 2016 Chinese Society of Aeronautics and Astronautics. Production and hosting by Elsevier Ltd.

Numerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant

Multi-phase flowfield simulation has been performed on solid rocket motor and effect of multi-phases on the performance prediction of the solid rocket motor(SRM) is investigation.During the combustion of aluminized propellant,the aluminum particles in the propellant melt and formliquid aluminum at the burning propellant surface.So the flow within the rocket motor is multi phase or two phase because it contains droplets and smoke particles of Al2O3.Flowsi mulations have been performed on a large scale motor,to observe the effect of the flowfield onthe chamber and nozzle as well.Uniform particles diameters and Rosin-Rammler diameter distribution method that is based on the assumption that an exponential relationship exists betweenthe droplet diameter,dand mass fraction of droplets with diameter greater thandhave been used for the si mulation of different distribution of Al2O3 droplets present in SRM.Particles sizes in the range of 1-100μm are used,as being the most common droplets.In this approachthe complete range of particle sizes is dividedinto a set of discrete size ranges,eachto be defined by single streamthat is part of the group.Roe scheme-flux differencing splitting based on approxi mate Riemann problem has been used to si mulate the effects of the multi-phase flowfeild.This is second order upwind scheme in which flux differencing splitting method is employed.To cater for the turbulence effect,Spalart-All maras model has been used.The results obtained show the great sensitivity of this diameters distribution and particles concentrations to the SRMflowdynamics,primarily at the motor chamber and nozzle exit.The results are shown with various sizes of the particles concentrations and geometrical configurations including models for SRM and nozzle.The analysis also provides effect of multi-phase on performance prediction of solid rocket motor.

Uncertainty analysis and probabilistic design optimization of hybrid rocket motors for manned lunar landing

To obtain a conceptual design for a hybrid rocket motor(HRM)to be used as the Ascent Propulsion System in the Apollo lunar module,the deterministic design optimization(DDO)method is applied to the HRM design.Based on the results of an uncertainty analysis of HRMs,an uncertainty-based design optimization(UDO)method is also adopted to improve the design reliability.The HRM design process,which is a multidisciplinary system,is analyzed,and a mathematical model for the system design is established to compute the motor performance from the input parameters,including the input variables and model parameters.The input parameter uncertainties are quantified,and a sensitivity analysis of the model parameter uncertainties is conducted to identify the most important model parameter uncertainties for HRMs.The DDO and probabilistic UDO methods are applied to conceptual designs for an HRM to be used as a substitute for the liquid rocket motor(LRM)of the Ascent Propulsion System.The conceptual design results show that HRMs have several advantages as an alternative to the LRM of the Ascent Propulsion System,including nontoxic propellant combination,small motor volume,and comparable functions,such as restarting and throating.Comparisons of the DDO and UDO results indicate that the UDO method achieves more robust and reliable optimal designs than the DDO method.The probabilistic UDO method can be used to develop better conceptual designs for HRMs.

Hybrid uncertainty-based design optimization and its application to hybrid rocket motors for manned lunar landing

Design reliability and robustness are getting increasingly important for the general design of aerospace systems with many inherently uncertain design parameters.This paper presents a hybrid uncertainty-based design optimization（UDO） method developed from probability theory and interval theory.Most of the uncertain design parameters which have sufficient information or experimental data are classified as random variables using probability theory,while the others are defined as interval variables with interval theory.Then a hybrid uncertainty analysis method based on Monte Carlo simulation and Taylor series interval analysis is developed to obtain the uncertainty propagation from the design parameters to system responses.Three design optimization strategies,including deterministic design optimization（DDO）,probabilistic UDO and hybrid UDO,are applied to the conceptual design of a hybrid rocket motor（HRM） used as the ascent propulsion system in Apollo lunar module.By comparison,the hybrid UDO is a feasible method and can be effectively applied to the general design of aerospace systems.

Steady-state coupled analysis of flowfields and thermochemical erosion of C/C nozzles in hybrid rocket motors

A hybrid rocket can be used in various applications and is an attractive propulsion system. However, serious erosion of nozzles is common in motor firing operations, which could restrict the application of hybrid rocket motors. Usually, the serious erosion is attributed to the high concentration of oxidizing species in hybrid motors, while the details of flowfields in the motors are not paid special attention to. In this paper, first the thermochemical erosion of C/C nozzle is simulated coupled with the flowfields in a 98% H2O2/hydroxyl-terminated polybutadiene（HTPB） hybrid rocket motor. The simulation is made on a typical axisymmetric motor, including a pre-combustion chamber, an aft-combustion chamber and nozzle structures. Thermochemica reactions of H2 O, CO2, OH, O and O2 with C are taken into account. Second, the change of flowfields due to fuel regression during motor firing operations is considered. Nozzle erosion in different flowfields is evaluated. Third, the results of nozzle erosion in the coupled simulation are compared with those under uniform and chemical equilibrium flow and motor firing test results. The results of simulation and firing tests indicate that the thermochemical erosion of nozzles in hybrid motors should be calculated coupled with flowfields in the motor. In uniform and chemical equilibrium flowfields, the erosion rate is overestimated. The diffusion flame in hybrid motors protects the nozzle surface from the injected oxidizer and high temperature products in flowfields, leading to a relatively fuel-rich environment above the nozzle. The influence of OH and the geometry of motor should also be considered in the evaluation of nozzle erosion in hybrid motors.

Geometric effects of fuel regression rate in hybrid rocket motors

The geometric configuration of the solid fuel is a key parameter affecting the fuel regression rate in hybrid rocket motors. In this paper, a semi-empirical regression rate model is developed to investigate the geometric effect on the fuel regression rate by incorporating the hydraulic diameter into the classical model. The semi-empirical model indicates that the fuel regression rate decreases with increasing hydraulic diameter and is proportional to dh?0.2 when convective heat transfer is dominant. Then a numerical model considering turbulence, combustion, solid fuel pyrolysis, and a solid–gas coupling model is established to further investigate the geometric effect. Eight motors with different solid fuel grains are simulated, and four methods of scaling the regression rate between different solid fuel grains are compared. The results indicate that the solid fuel regression rates are approximate the same when the hydraulic diameters are equal. The numerical results verify the accuracy of the semi-empirical model.

Effect of the head cavity on pressure oscillation suppression characteristics in large solid rocket motors

In order to discover the effect of head cavity on resonance damping characteristics in solid rocket motors, large-eddy simulations with wall-adapting-local-eddy-viscosity subgrid scale turbulent model are implemented to study the oscillation flow field induced by vortex shedding based on the VKI （yon Karman Institute） experimental motor. Firstly, mesh sensitivity analysis and grid-independent analysis are carried out for the computer code validation. Then, the numerical method is further validated by comparing the calculated results and experimental data. Thirdly, the effects of head-end cavity on the pressure oscillation am-plitudes are studied in this paper. The results indicate that cavity volume, location and configuration have a cooperative ef- fect on the oscillation amplitude. It is proved that Rayleigh criterion can be used as a guiding principle for the design of reso- nance damping cavity. The change of the head-end cavity breaks the balance between the mass flux and acoustic energy. Therefore, the pressure oscillation characteristics change accordingly. It is concluded that a large mass flux added at the pres- sure antinode could attribute to significant amplitude. Meanwhile, the damping effect of the cavity is stronger when the dis- tance between cavity and pressure antinode becomes shorter. Finally, this method is applied to the modification of an engi- neering solid rocket motor. The static test of solid rocket motor reflects that the oscillations can be effectively suppressed by a head-end cavity.

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.

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.

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.

Center of rotation estimation for rocket nozzle by infrared reflective makers and image sequences

The determination of an accurate center of rotation of rocket motor nozzle or other object to be measured is of great interest across a wide range of applications,such as rocket,missile,robotics,industry,spaceflight,aviation and human motion analysis fields,particularly for clinical gait analysis.A new approach was proposed to estimate the moving objects' instantaneous center of rotation and other motion parameters.The new method assumes that the two segment of object to be measured are rigid body which rotates around a center of rotation between each other relatively.The center of rotation varies with time in the global coordinate system but is fixed in the local coordinate system attached to each segment.The models of rocket motor nozzle and its movement were established.The arbitrary moving object's corresponding to motion equations were deduced,and the least square closed-form solutions of the object's motion parameters were figured out.It is assumed that the two high speed CCD cameras mounted on the 750 nm infrared(IR) filter are synchronized and calibrated in advance.The virtual simulation experiment using 3D coordinates of markers was conducted by synchronized stereo image sequences based on 6-DOF motion platform and the experimental results prove the feasibility of our algorithm.The test results show that the precision of x,y,z component on center of rotation is up to 0.14 mm,0.13 mm,0.15 mm.